The Development of a Pulsed Plasma Thruster as a Solid Fuel Plasma Source for a High Power Helicon

Ian Kronheim Johnson

A thesis submitted in partial fulfillment of the requirements for the degree of

Master of Science

The University of Washington

2011

Program authorized to offer degree: Aeronautics and Astronautics

University of Washington

Graduate School

This is to certify that I have examined this copy of a master’s thesis by

Ian Kronheim Johnson

and have found that is it complete and satisfactory in all respects, and that any and all revisions required by the final examining committee have been made.

Committee Members:

Professor Robert Winglee, Department of Earth and Space Sciences, Chair

Professor Tom Jarboe, Department of Aeronautics and Astronautics

Date:

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In presenting this thesis in partial fulfillment of the requirements for a master’s degree at the University of Washington, I agree that the Library shall make its copies freely available for inspection. I further agree that extensive copying of this thesis is allowable only for scholarly purposes, consistent with “fair use” as prescribed in the U.S. Copyright Law. Any other reproduction for any purposes or by any means shall not be allowed without my written permission.

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Abstract

The Development of a Pulsed Plasma Thruster as a Solid Fuel Plasma Source for a High Power Helicon

Ian Kronheim Johnson

Chair of the Supervisory Committee:

Professor Robert Winglee

Earth and Space Sciences

As space exploration shifts to lower mass and lower cost missions, the need for improved on-board propulsion systems is growing. The High Power Helicon (HPH) experiment at the University of Washington is one new possibility in this field. The HPH is currently gas-fed, which creates a number of added concerns during operation. For this reason, a Pulsed Plasma Thruster (PPT) was developed to act as a solid fuel source for the Helicon thruster. The PPT created was 50% smaller in size than the EO-1 satellite PPT [Zakrzwski] and produced electron densities on the order of 1018 m-3, 20 cm downstream of the Teflon surface at 43 J. The thruster had an estimated impulse bit of 138 µN-sec and a specific impulse of 1413 seconds while ablating 44 µg of Teflon per pulse at 4.4% efficiency. The HPH system currently uses a helicon antenna to ionize and accelerate the gas particles with high efficiency. Therefore, the objective of the PPT for the HPH operating system was to produce large numbers of neutral particles traveling at low to medium velocities. This is in comparison to most PPT investigations where higher efficiencies and thrusts are desired. The PPT was successfully fired under vacuum while operated with the HPH magnetic field and is ready to be fully integrated and tested with the HPH system.

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Table of Contents

1 Introduction...... 1 1.1 Current Propulsion Concepts ...... 1 1.1.1 Electrostatic Propulsion...... 2 1.1.2 Electromagnetic Propulsion...... 3 1.1.3 Electrothermal Propulsion...... 5 1.1.4 Chemical and Electrical Comparison...... 7 1.2 High Power Helicon Thruster...... 8 1.3 Helicon Thruster Design...... 10 1.3.1 Gas vs. Solid Fed Helicon Thruster...... 11 1.4 Pulsed Plasma Thruster...... 12 1.4.1 PPT Operation ...... 13 1.4.2 PPT Plume ...... 15 1.4.3 PPT Flight History...... 16 1.5 Outline and Research Goals...... 19 2 Experimental Setup and Electronics Design...... 21 2.1 Vacuum System ...... 21 2.2 Control and Data Acquisition ...... 22 2.3 Electronic Control Systems ...... 23 2.3.1 Starter Power Processing Unit ...... 23 2.3.2 Main Discharge Power Processing Unit ...... 24 2.4 HPH Base Magnets ...... 27 3 Diagnostics...... 28 3.1 Stangenes Probe...... 28 3.2 Voltage Dividers and the High Voltage Probe ...... 28 3.3 Langmuir Probes...... 29 3.3.1 Standard Double Probe Theory ...... 30 3.3.2 Probe Construction and Description ...... 34 3.4 CCD Camera...... 35 4 Theoretical Models and Data Reduction ...... 36 5 PPT Iterations and Discussion...... 38 5.1 PPT-1...... 38 5.1.1 Spark Plug – PPTv1.0...... 38

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5.1.2 Spark Wire – PPTv1.1 & 1.2...... 39 5.2 PPT - 2...... 41 5.2.1 No Insulation - PPTv2.0...... 41 5.2.2 Insulation - PPTv2.1...... 42 5.2.3 Electron Density On-axis...... 43 5.2.4 Plume Width...... 47 5.2.5 Predicted Electron Densities at 10 cm height...... 54 5.2.6 Time-of-Flight Velocity Data ...... 54 5.2.7 Mass Ablation Estimates...... 57 5.2.8 PPT Performance Estimates...... 58 5.2.9 Effect of an Applied Magnetic Field...... 59 6 Summary...... 63 7 Bibliography...... 65

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List of Figures

Figure 1.1. An electrostatic gridded ion thruster schematic (left) [XIPS] and a xenon ion discharge from the NSTAR ion thruster (right) [Ion Propulsion]...... 2 Figure 1.2. The MPD schematic (left) [Jahn and Choueiri] and (right) a NASA MPD thruster experiment in operation (right) [MPD: NASA Facts]...... 4 Figure 1.3. The Hall thruster schematic (left) [Jahn and Choueiri] and a NASA Hall thruster firing with xenon propellant (right) [Electric Propulsion]...... 5 Figure 1.4. A resistojet schematic (left) [Martinez-Sanchez and Pollard] and a flight ready hydrazine resistojet built by the Primex Corporation (right) [Jahn and Coueiri]...... 6 Figure 1.5. An arcjet schematic (left) [Martinez-Sanchez and Pollard] and an ammonia arcjet in operation (right) [Zube]...... 6 Figure 1.6. The VASIMR electromagnetic thruster schematic (top) and the VASIMR thruster firing in its first full power testing (bottom) [Ad Astra Rocket Company]...... 9 Figure 1.7. The right-handed helical antenna used in the High Powered Helicon thruster (left) and the Helmholtz style magnetic coil (right)...... 11 Figure 1.8. Schematic of a typical PPT with rectangular electrodes [Eckman]...... 13 Figure 1.9. The first PPT designs from the Soviet Union. The electromagnetic (right) and electrothermal (left) designs were both considered. The Zond-2 spacecraft used the electrothermal design [Burton and Turchi]...... 16 Figure 1.10. Shown are the Soviet Union’s Zond-2 Satellite and the American LES-6 satellite [Burton and Turchi]. Both utilized pulsed plasma thrusters as their primary propulsion system...... 17 Figure 1.11. The EO-1 spacecraft with attached pulsed plasma thruster (left). The PPT is attached to the right of the spacecraft. In the image on the right, a 12-inch ruler is shown to give scale [Zakrzwski]...... 18 Figure 1.12. The Dawgstar PPT in operation (left) [Rayburn] and the µPIT thruster nozzle (right) [Peters]...... 19

Figure 2.1. The bell jar vacuum chamber enclosed within a plexiglass shield. The mounting flange is bolted to the table and the glass jar is connected to the flange only through the pressure difference imposed upon the o-ring...... 21 Figure 2.2. Shown here are the turbopump, roughing pump, and all three gauges connected to the chamber. This setup lies directly below the bell jar...... 22 Figure 2.3. The electronic schematic for the PPT showing both the starter arc and the main discharge...... 23

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Figure 2.4. The circuitry for the starter arc which is housed within the electronics box with the main discharge circuitry...... 24 Figure 2.5. The main discharge circuitry is shown here, including the 60 μF capacitor, 15 kV diodes and the dump board...... 25 Figure 2.6. Generic voltage traces for the PPT main discharge without the diode (left) and with the diode (right). The traces are shown for four different energy input levels (120 J in blue, 43 J in green, 20 J in red, and 10 J in turquoise)...... 26 Figure 2.7. Generic current traces for the PPT main discharge without the diode (left) and with the diode (right). The traces are shown for four different energy input levels (120 J in blue, 43 J in green, 20 J in red, and 10 J in turquoise)...... 26 Figure 2.8. Base magnets over the PPT in the bell jar while at atmospheric pressure. A Langmuir probe can be seen in the right image 20 cm above the Teflon surface. . 27

Figure 3.1. Voltage divider circuit diagram. For this application and ...... 28

Figure 3.2. The circuit diagram for a double Langmuir probe...... 30 Figure 3.3. Double Langmuir probe potential diagram [Byrne]...... 30 Figure 3.4. Theoretical IV characteristic for a symmetric double Langmuir probe... 34 Figure 3.5. The symmetric double Langmuir probe used for these experiments...... 34

Figure 5.1. The first iteration of the PPT with the spark plug imbedded within the cathode...... 38 Figure 5.2. The PPT with a spark wire (v1.1) is shown here at 20 µsec into the pulse (top right), 60 µsec (lower left), and 100 µsec (lower right). The PPT was fired with an input voltage of 300 V...... 39 Figure 5.3. PPTv1.2 is shown here with torr seal added to the back and sides of both electrodes and Kapton tape added over the high voltage screw connected to the anode. The images are shown at input voltages of 200 V (top right), 300 V (lower left), and 400 V (lower right)...... 40 Figure 5.4. PPT 2.0 is shown at 20 µsec into the pulse (top left), 40 µsec (top right), 60 µsec (bottom left), and 80 µsec (bottom right) with a voltage of 400 V...... 41 Figure 5.5. PPT2.1 is shown at 200 V (1.2 J) and 40 µsec into the shot. The torr seal was placed over the anode only – including the connection point where the high voltage lead was soldered to the electrode...... 42 Figure 5.6. End on views of the PPT firing at 1200 V (43 J) with an exposure covering the entire length of the shot...... 42

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Figure 5.7. An example of the sputtering affect on Langmuir probes at 43 J. The blue trace was taken at a battery voltage of 120 V, while the red trace was taken at 129 V. In this case the maximum density level and time could be recorded, however if the sputtering effect had occurred before the peak, then the data would be useless...... 43 Figure 5.8. Typical, non-sputtering Langmuir traces taken at 10, 20, and 43J at 20cm as a function of time. These three traces correspond to the dark blue traces in Figures 5.9 – 5.11...... 44 Figure 5.9. Electron density and temperature results on-axis at 10J...... 45 Figure 5.10. Electron density and temperature results on-axis at 20J...... 45 Figure 5.11. Electron density and temperature results on-axis at 43J...... 46 Figure 5.12. Peak electron density on-axis and at 6 and 8 cm off-axis at 10 J at heights between 20 and 35 cm above the Teflon surface...... 48 Figure 5.13. Peak electron density on-axis and at 6 and 8 cm off-axis at 20 J at heights between 20 and 40 cm above the Teflon surface...... 49 Figure 5.14. Peak electron density on-axis and at 6 and 8 cm off-axis at 43 J at heights between 20 and 40 cm above the Teflon surface...... 50 Figure 5.15. The Gaussian Approximations Full-Width Half-Maximum values for 10 (green), 20 (red) and 43 J (blue) between 20, 25, 30, 35, and 40 cm above the Teflon surface. Opening angles of 47.5, 49.2, and 58.6° were found for the 10, 20, and 43 J cases by taking the tangent of the slopes of the lines of best fit...... 52 Figure 5.16. The electron area density calculated values based on Eq. 5.5 for 10, 20, and 43 J taken between 20 and 40 cm...... 53 Figure 5.17. TOF velocity data from the electron density results at 10 J...... 55 Figure 5.18. TOF velocity data from the electron density results at 20 J...... 55 Figure 5.19. TOF velocity data from the electron density results at 43 J...... 56 Figure 5.20. The PPT in operation with the helicon magnets producing a magnetic field of approximately 300 G...... 59 Figure 5.21. The PPT in operation within a 300 G field and a 7 cm quartz tube placed over the thruster exit. The image on the left was taken with a 20µsec exposure length and the image of the right with a 60 µsec length...... 60 Figure 5.22. Electron density results on-axis in the presence of a 300 G applied magnetic field at 10J and 30cm above the Teflon surface...... 61 Figure 5.23. Electron density results on-axis in the presence of a 300 G applied magnetic field at 20J and 30cm above the Teflon surface...... 61 Figure 5.24. Electron density results on-axis in the presence of a 300 G applied magnetic field at 43J and 30cm above the Teflon surface...... 62

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List of Tables

Table 1.1. A comparison of the six general chemical and electrical propulsion concepts. Notice that these propulsion systems span four orders of magnitude in specific impulse and twelve orders of magnitude in thrust [Sutton]...... 7 Table 1.2. The thrust and impulse characteristics of the pulsed plasma thrusters discussed previously. All the listed thrusters have a firing rate of 1 Hz. For variable thrust capability, the maximum values were listed [Burton and Turchi unless otherwise noted]. * signifies the PPT was never flown...... 19

Table 5.1. The electron density and temperature Langmuir data on-axis for the PPT at 10, 20, and 43 J and heights from 20 to 40 cm in 5 cm increments...... 46 Table 5.2. Constants in Eq. 5.2 used for the 10 J Gaussian density approximations in Figure 5.12...... 48 Table 5.3. Constants in Eq. 5.2 used for the 20 J Gaussian density approximations in Figure 5.13...... 49 Table 5.4. Constants in Eq. 5.2 used for the 43 J Gaussian density approximations in Figure 5.14...... 50 Table 5.5. The on-axis and off-axis electron density measurements taken at 10, 20, and 43 J. Off-axis measurements were made at 6 and 8 cm from the Teflon edge..... 51 Table 5.6. The peak electron area density average values and their corresponding uncertainties...... 53 Table 5.7. The off-axis electron density measurements taken at a height of 10 cm at 10, 20, and 43 J energy levels...... 54 Table 5.8. A summary of the TOF velocity measurements made at 10, 20, and 43 J in km/s and eV. Uncertainty in Langmuir measurements give errorbars of ±6 km/sec and ±14 eV...... 56 Table 5.9. Comparison of the ablated mass per pulse for the PPT at 20, 30, and 40 cm heights for the 10, 20, and 43 J pulse energies...... 57 Table 5.10. Comparison of the ablated mass per pulse for the LES8/9 PPT [Vondra, et. al], µPPT at the Air Force Research Laboratory [Spanjers, 2001], and an electrothermal pulsed plasma thruster currently being tested at the University of Illinois [Bushman and Burton]...... 58

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Acknowledgements

I wish to express my profound gratitude to my advisor, Prof. Robert Winglee, for his extensive assistance throughout my research. Especially these past few weeks as he put up with my procrastination in regards to the actual writing of the thesis. I look forward to working alongside him over the upcoming years on my doctorate.

Even though Dr. Tim Ziemba was managing multiple projects at EHT, he always had time to answer my questions and give advice on the solid fuel PPT.

I want to thank Prof. Tom Jarboe for serving on my committee and all the insight he gave throughout the courses i've taken from him over the past two years.

This project never would have succeeded without the help of Dr. Jim Prager, Race Roberson, and Ilia Slobodov here in the APL. Their assistance, guidance, and opinions helped get this project started and running.

I would never have been here without the patience, assistance, and guidance of my parents, Orlay and Shirley, and my sister Shannon. Thank you for everything over the past twenty five years. Good luck in Chicago Shannon.

And finally to all those friends who helped take my mind off research. Whether it was a drink on the Ave after school, a Husky Football game on the weekend, or a soccer game on a rainy and cold weeknight - thank you for being there.

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1 Introduction

Higher performance and longer lifetimes of propulsion systems allow mission designers to provide more science payload capability and higher reliability to their spacecraft. The need for low-cost and high payload mass fractions for in-space thrusters is currently a driving force for new propulsion research within NASA and around the world for deep space and long lifetime Earth-orbital missions [Dunning]. Traditional chemical propulsion thrusters cannot perform these missions economically or within reasonable time frames. For example, Voyager I and II, launched in 1977 use sixteen monopropellant thrusters apiece and to date they have yet to leave our solar system [Angrum]. Current spacecraft sent to the outer planets using advanced chemical thrusters still have mission times on the order of tens of years and require gravitation assists from other planets - severely limiting launch windows. Long trip times and lower payload fractions limit the scientific gain of these missions. Even typical missions, such as long duration Earth orbits or lunar expeditions, need lighter propulsion systems in order to reduce the initial launch weight and correspondingly the mission cost. Therefore, new technology must be developed and produced to enable the mission scenarios being considered today. 1.1 Current Propulsion Concepts

Modern space propulsion techniques can be separated into two categories: chemical and electrical systems. Chemical systems use the energy stored in the bonds of the propellant to produce a high temperature, resulting in a high pressure working fluid that can then be expelled from a conventional, converging-diverging nozzle to produce thrust. The fluid is nearly always a gas created by a high pressure combustion of solid or liquid propellants consisting of fuel and oxidizer components within a combustion chamber [Peters]. The main areas of research within this field are solid propellants, liquid-bipropellants, monopropellants, and cold-gas thrusters. All of these engines have been successfully used on multiple missions ranging from the drag make-up of satellites to the main propulsion system on interplanetary science probes [Sutton]. These technologies have been thoroughly developed over the past half-century and reap the benefits of that heritage in the consideration for placement on current missions.

The second category, electric propulsion, uses electrical energy to produce an ionized gas which can then be accelerated from the engine using one of three main techniques.

Electrostatic – The acceleration is caused predominantly by the Coulomb force, where a static electric field in the direction of the acceleration is applied.

Electromagnetic – Orthogonal electric and magnetic fields apply a Lorentz body force to highly ionized propellant atoms, accelerating them out of the thruster.

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Electrothermal – The propellant gas is heated electrically and then expanded in a nozzle. Similar to chemical propulsion except the heating is done electrically and not chemically.

1.1.1 Electrostatic Propulsion

Electrostatic thrusters rely on Coulomb forces to accelerate a propellant composed of non-neutral charged particles. The electric force depends only on the charge, and all charged particles must be the same “sign” if they are to move in the same direction [Sutton]. Electrons are easy to produce and are readily accelerated, but they are so extremely light in mass that they are impractical for propulsion as their momentum is small, even at velocities nearing the speed of light. Accordingly, electrostatic thrusters use charged heavy-molecular-mass atoms as positive ions. A proton is 1836 times heavier than an electron and a typical ion of interest contains hundreds of protons [Sutton]. The most common electrostatic device is the gridded ion thruster. As Figure 1.1 shows, in a gridded ion thruster the gas is pumped from the back and accelerated forward by the potential difference between the negative electron grid and the plasma itself.

Figure 1.1. An electrostatic gridded ion thruster schematic (left) [XIPS] and a xenon ion discharge from the NSTAR ion thruster (right) [Ion Propulsion].

Currently, a gridded ion thruster is in use on the Dawn satellite, a NASA spacecraft sent to two of the largest members of the asteroid belt – Vesta and the dwarf planet Ceres. It was launched in 2007 and reached Vesta on July 16th, 2011, which it will orbit and explore until 2012 [Brophy]. Dawn is scheduled to reach Ceres in 2015 and will be the first spacecraft to visit either body. The mission is intended to answer questions about the formation of the solar system. Vesta and Ceres were chosen as two contrasting protoplanets, the first “dry”, or rocky, and the second “wet”, or icy [Russell]. The ion propulsion system includes three 30-cm diameter xenon ion thrusters, built in a similar design to the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR). NSTAR was the first time an ion engine was used as

The University of Washington 2 the primary propulsion system for a deep space mission [Brophy]. The Dawn spacecraft began with 450 kg of xenon and is designed to operate over a 10-year time period. The Dawn thrusters can provide a total of 11 km/s [Brophy].

1.1.2 Electromagnetic Propulsion

In contrast to an electrostatic device, an electromagnetic thruster produces an electrical current within a conductive propellant in the presence of a magnetic field. The magnetic field interacts with the current to generate a force on the propellant, expelling it out of the thruster [Jahn and Choueiri]. This force is referred to as the Lorentz force. Electrostatic systems generally have higher efficiencies than electromagnetic thrusters, which mean they need a lower input power, but they have relatively low number and energy densities so that high power configurations are unfortunately rather large [Sutton]. Electromagnetic thrusters are characterized by higher energy densities and more compact designs [Sutton].

Unlike electrothermal and electrostatic devices, which offer only a few practical configurations, electromagnetic acceleration presents multiple possibilities for implementation. They have potential usage from microsatellites at <1 W [Peters] to interplanetary human exploration missions at >200 MW of power [Ilin]. The applied fields and internal currents can be steady, pulsed, or alternating over a broad range of frequencies; the magnetic fields may be externally applied or induced by the current patterns; and a broad range of propellant types have been employed [Jahn and Choueiri]. The most widely used electromagnetic thrusters are magnetoplasadynamic (MPD) thrusters, Hall-current accelerators, and pulsed plasma thrusters (the pulsed plasma thruster is discussed in Section 1.4).

The MPD thruster is characterized by a coaxial geometry with inner and outer electrodes as well as an inter-electrode insulator. As shown in Figure 1.2 gaseous propellants are introduced into the upstream portion of the channel, where after they are ionized by passage through an intense electric arc standing in the inter- electrode gap [Jahn and Choueiri]. The cross product between this arc current and its associated magnetic field (Loretnz force) directs the propellant downstream creating an extremely hot plasma just beyond the cathode tip. The nozzle configuration dictates the expansion of plasma to yield the requisite exhaust velocity [Jahn and Choueiri].

Although MPD thruster technology has been explored on the ground, commercial interest has been low due to two core problems. Optimum performance of the MPD requires powers on the order of hundreds of kilowatts and current interplanetary spacecraft power systems are incapable of producing those power levels [LaPointe and Mikellides]. NASA’s Project Prometheus nuclear reactor was expected to generate power in the hundreds of kilowatts range, but was discontinued in 2005 [Oleson]. The second main concern is the degradation of the cathode due to evaporation of the copper due to high current densities. Experimentation at the University of Southern California has shown that barium and lithium propellant

The University of Washington 3 mixtures and multi-channel hollow cathodes are a possible solution to the erosion problem [Downey].

Figure 1.2. The MPD schematic (left) [Jahn and Choueiri] and (right) a NASA MPD thruster experiment in operation (right) [MPD: NASA Facts].

Hall thrusters on the other hand have been flown in orbit since 1971 when the Soviet Union launched the SPT-50 thruster on board the Meteor spacecraft [Sutton]. Over 200 Hall thrusters have flown over the past forty years without a single failure in orbit [Sutton]. The first use of Hall thrusters outside of Earth’s orbit was on the European Space Agency (ESA) lunar mission SMART-1 in 2003. The Hall thruster performed flawlessly until 2006 when the spacecraft was deliberately crashed into the Moon’s surface [Koppel and Estublier].

Similar to the MPD thruster, the Hall thruster has a coaxial geometry and an emissive cathode surface. The cathode emission produces an attractive negative charge downstream from the gas injection point. This negative charge acts similar to the negative grid in the electrostatic thruster in that it creates an electrostatic potential to accelerate the ions up to high speeds [Jahn and Choueiri].

In the simplest picture, the applied electric field accelerates ions and the magnetic field prevents electrons from shorting out the voltage. The azimuthal current produced applies a force to the rocket and thus, this is a Lorentz force accelerator and not an electrostatic thruster as often thought [Goebel and Katz]. Hall thrusters have optimum performance at considerably lower power levels than MPD devices due to the low mass flow densities and the fact that the magnetic fields are externally produced [Jahn and Choueiri]. The Hall thruster can be seen firing in Figure 1.3.

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Figure 1.3. The Hall thruster schematic (left) [Jahn and Choueiri] and a NASA Hall thruster firing with xenon propellant (right) [Electric Propulsion].

1.1.3 Electrothermal Propulsion

Unlike conventional rocket engines that produce thrust by means of complex chemical reactions, electrothermal devices use strong electric current to directly heat and accelerate an inert propellant gas [Harris]. This is the closest electric propulsion method to the classical chemical rocket engine. Electrothermal propulsion comprises all methods where a propellant is electrically heated in a chamber and then expanded through a nozzle to convert its thermal energy to a directed stream which delivers thrust to the satellite [Jahn and Chouriei]. Common electrothermal thrusters are the resistojet and arcjet.

Resistojets transfer heat from a solid surface such as a heater coil or the chamber wall to the propellant. Figure 1.4 shows a heater coil resistojet design. The goal is to impart additional energy to the plasma, thus extracting the maximum energy per kilogram of the propellant, at the expense of increased power consumption [Jahn and Coueiri]. However, the wall or coil has an upper limit to the temperature at which it can withstand, creating an upper boundary to the energy which can be imparted to the propellant. This physical temperature limitation to the coil is the primary limitation to the resistojet [Sutton]. The arcjet attempts to overcome this concern by removing the coil from the system altogether and replacing it with an electric arc.

The first resistojet was flown onboard the American satellite Vela in 1965 for orbit phase adjustment. It wasn’t until 1999 that a resistojet was launched by a European country, flying onboard the Survey Satellite Technology (SSTL) UoSAT 12 satellite by the United Kingdom for orbit correction and adjustment [Martinez-Sanchez]. Over the past decade arcjets have been favored over the resistojet as the temperature limitations are not as drastic and higher performance can be achieved.

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Figure 1.4. A resistojet schematic (left) [Martinez-Sanchez and Pollard] and a flight ready hydrazine resistojet built by the Primex Corporation (right) [Jahn and Coueiri].

In the arcjet, an electrical discharge (an arc) is created in the flow of the propellant to impart additional energy to the plasma. Similarly to a resistojet, this is designed to extract the maximum energy per kilogram of the propellant, at the expense of increased power consumption. Unlike resistojets, the wall temperature limitation is overcome by depositing power internally [Walker]. The gas is typically injected with an azimuthal swirl which prevents the high temperature arc from bending and touching the walls [Sutton].

Numerous arcjets have been space qualified and are currently in use. The first commercial application of an arcjet occurred in 1993 with the launch of Telstar IV which used four 1.8 kW hydrazine arcjets for north-south station keeping [Walker]. The highest power arcjet to ever fly (as well as the highest power electric propulsion thruster to ever fly) is the 26 kW ammonia arcjet which flew as part of the electric Propulsion Space Experiment (ESEX). ESEX was launched in 1999 to study the effect of high power electric propulsion systems on other spacecraft functions [Bromaghim].

Figure 1.5. An arcjet schematic (left) [Martinez-Sanchez and Pollard] and an ammonia arcjet in operation (right) [Zube].

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1.1.4 Chemical and Electrical Comparison

As was previously mentioned, chemical systems have the advantage of being well tested in the laboratory and in space flight. As the propellants in a chemical system do not rely on outside energy sources, the rate at which energy can be supplied is independent of the mass of the propellant – resulting in extremely high power and thrust levels, as shown in Table 1.1. However, chemical propulsion is “energy limited”, in that the chemical reactants have a fixed amount of energy per unit mass [Electric Spacecraft Propulsion]. This limits the achievable exhaust velocity and specific impulse. Additionally, the mass of fuel within a chemical propulsion system becomes exponentially larger as the mission duration increases. This results in decreased burn times - severely handicapping the mission length and increasing mission timelines [Sutton].

Electric propulsion systems are said to be "power limited" as the rate at which energy from an external source is supplied to the propellant is limited by the mass available for the power system [Electric Spacecraft Propulsion]. This results in limiting the thrust of an electric propulsion system to the milli-newton range for moderate power levels. Because of this, electric propulsion vehicles tend to have low thrust-to-mass ratio (and correspondingly low possible accelerations). However, electric propulsion devices can have large specific impulses and exhaust velocities. Due to the low fuel mass utilized, these systems can run for years, compared to the minutes to hours for which chemical systems burn for. This means that the total impulse level (total momentum change) of an electric propulsion thruster can be much greater than for a chemical device of the same mass [Jahn and Choueiri].

The low thrust levels make electric propulsion devices poorly suited for launch vehicles or other mission aspects that require rapid changes in velocity. However, the long burn times and large total impulses make electric propulsion ideal for deep space or small precision thrust missions [Sutton].

Table 1.1. A comparison of the six general chemical and electrical propulsion concepts. Notice that these propulsion systems span four orders of magnitude in specific impulse and twelve orders of magnitude in thrust [Sutton].

Thruster Concept Exhaust Velocity Isp (sec) Thrust (km/s) (mN) Monopropellant 1-3 125 - 250 100-104 Solid propellant 1-4 250 - 300 106 - 1010 Bipropellant 1.5-4.2 200 - 450 100 - 1010 Electrothermal 4-16 300 -1000 200 - 1000 Electromagnetic 10-300 800 - 5000 0.05-10 Electrostatic 15-210 2000 - 20000 0.01-200

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1.2 High Power Helicon Thruster

Since their introduction in the 1980’s by Perry and Boswell, helicon prodced plasma have been an active area of research [Perry, Boswell]. With the ability to produce an order of magnitude increase in plasma densities over standard inductive methods, they have become the plasma source of choice for multiple applications. These include possible heating mechanisms for fusion devices [Tripathi], space propulsion thrusters [Bowsell and Chen], and plasma material processing and etching [Aachboun and Ranson].

Helicon waves are right-handed, circularly-polarized, electromagnetic (whistler) waves that only propagate in radially confined magnetized media. Helicon plasma sources are capable of producing extremely high density plasma with high ionization efficiencies and as such may be a reasonable in-space thruster concept for future missions [Prager].

A whistler wave is a low frequency, circularly polarized, electromagnetic wave that propagates through a plasma parallel to the magnetic field. The first reports of whistler waves date back to World War I, where soldiers heard static that changed frequency from high to low over phones lines near the battle front. They described it as the whistling sound a grenade makes and called the phenomena “the grenade fly” [Boswell and Chen]. In 1919, Barkhausen first recorded these tones and later revisited these measurements in 1930 but was unsure how to explain them. [Barkhausen]

In 1953 Storey was able to calculate the maximum angle for the group velocity vector of whistler waves, which showed that these waves propagated within a cone around magnetic field lines. He also calculated a maximum group velocity, which meant that higher and lower frequency modes would reach the receiver later than modes traveling at the maximum group velocity, which would explain the sound the soldiers heard. He further concluded that lightning strikes generated a broad frequency spectrum and these waves would propagate through the atmosphere as whistler waves. [Storey]

In 1960, Aigrain produced the first laboratory helicon wave during his study of waves in solid metals [Chen]. He observed waves in slabs of super low-temperature sodium that propagated in the range of frequencies , where and are the ion and electron cyclotron frequencies, respectively. Upon further study, he determined that the wave magnetic field vector traced a helix at a fixed time, hence the name “helicon”.

Modern helicon plasmas are produced in cylindrical configurations with a DC magnetic field applied along the longitudinal axis. The gas is first weakly ionized by the electrostatic fields in the antenna region as in a typical capacitively (CCP) or inductively coupled plasma (ICP). However, upon application of the external magnetic field, the plasma discharge changes character in that it is no longer subject

The University of Washington 8 to the skin depth constraint, which the CCP and ICP are. This allows the helicon wave to penetrate into the core of the plasma column. The plasma is then further ionized due to a wave-particle interaction and is thought to be aided by a mode conversion at the wall boundary. [Pucci]

Today, helicon wave sources are being used for a variety of applications due to their ability to efficiently produce a uniform, high density plasma. For example, Chen has produced helicon plasmas with densities up to with uniformities of for use in materials processing devices [Chen]. A helicon source is being used as the primary ionization source in the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) concept (Figure 1.6) is currently being developed at the Ad Astra Rocket Company [Chang-Diaz]. Engineers have been able to routinely produce hydrogen, deuterium, and helium plasmas with peak densities of with the VASIMR helicon source [Jacobson]. The VASIMR system, operating at 12 MW, has been predicted to have a 3-4 month Earth-Martian transit time; less than half the time a nuclear or chemical rocket would take [Ilin].

Figure 1.6. The VASIMR electromagnetic thruster schematic (top) and the VASIMR thruster firing in its first full power test (bottom) [Ad Astra Rocket Company].

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Helicon sources have many advantages over conventional plasma sources. At the forefront of those advantages is that they are one of the most efficient laboratory sources of plasma currently known. The reason for their high efficiency is unknown; however it could be related to the mechanism by which the wave energy is transferred to the plasma [Prager]. As previously stated, they also have the ability to produce relatively dense plasmas in the range of . Another important advantage, although common to all RF plasma sources, is the fact that there are no electrodes in contact with the plasma, thus eliminating the possibility of contamination via sputtering [Prager]. 1.3 Helicon Thruster Design

The Helicon thruster has a simple design: a quartz tube wrapped by a coiled antenna, surrounded by magnets. The gas used as propellant is pumped into the quartz tube, where it is turned into plasma. The magnets confine, guide, and accelerate the plasma into an exhaust beam, which creates the thrust. In other helicon experiments, the antenna is wrapped around the outside of a dielectric vacuum chamber, and the magnets sit outside the antenna [Boswell and Vender]. Unlike previous experiments, HPH is housed inside the vacuum chamber so that it can be tested in a space-like environment, which keeps the helicon source far from the walls.

The antenna used in the HPH thruster is a left-handed, half wavelength, helical, Nagoya Type III antenna that has been adopted by Chen and many others [Chen and Boswell]. This antenna was chosen based on the work by Light and Chen [Light], who investigated plasma properties produced by an untwisted (N), right hand (RH) twisted, and left hand (LH) twisted Nagoya Type III antennas. The handedness of that antenna is defined with respect to the direction that the base magnetic field points. For example, if you move along the magnetic field of a RH antenna, then you will see the coil rotate clockwise around the field; and respectively rotate counterclockwise for a LH antenna. This antenna design was constructed to excite the helicon mode

The helicon antenna (shown on the left in Figure 1.7) was wrapped around a quartz tube with a diameter of 7 cm and a length of 15 cm. The antenna was constructed with flat braided wire 5.2 mm in width. The braided wire was chosen for its flexibility, making it easier to wind in a helical form. One end of the tube was capped with a circular quartz plate covered with boron nitrite, while the other end is open to allow plasma to escape downstream. Note that in Figure 1.7 there are two helical antennas wrapped around the quartz tube. These two antennas are rotated 90 degrees with respect to each other. The second antenna was designed to create a field that rotated in space as well as time, similar to the bifilar antennas used by [Miljak and Chen].

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Figure 1.7. The right-handed helical antenna used in the High Powered Helicon thruster (left) and the Helmholtz style magnetic coil (right).

The HPH thruster uses a Helmholtz style coil (Figure 1.7 right), which produces a magnetic field that is nearly uniform between the coils. The magnetic coils are typically operated with all six magnets energized to create a solenoid magnetic field that points downstream with 200-500 Gauss on the center axis.

The thruster is housed in a vacuum chamber 2.8 meters long. 0.8 meters in radius, and capable of reaching pressures of 1×10-6 Torr. This large size and low pressure allows not only for the study of the helicon as a thruster, but also the physics behind Helicon waves.

1.3.1 Gas vs. Solid Fed Helicon Thruster

As the gas is pumped from exterior tanks into the quartz tube, the magnets confine, guide, and accelerate the plasma into an exhaust beam - which creates the thrust. However, there are a number of concerns while using a gas-fed system such as this. First and foremost the storage of a gas into space adds significantly to the mass of the system. Feeding the thruster with gas requires a complex system of tubes, valves, and tanks that are all under an extremely high pressure. Such a complex system increases the mass of the thruster and the amount of maintenance. It is possible to use liquids that can then be vaporized but this again adds to the complexity of the system.

A secondary problem with gas feed systems is that gas flows with a low velocity compared to that of the plasma velocity, creating a discontinuity which prevents the system for being run in an optimal operating condition.

Replacing the gas input with a solid feed system would be one way to overcome these concerns. This solid feed system would need to produce a semi-continuous stream of neutral particles. One of the simplest and most effective solid feed systems is the Pulsed Plasma Thruster (PPT). Instead of pumping gas into the chamber, a PPT would already have the propellant within the thruster and, theoretically, be able to run continuously. One of the largest advantages to a PPT is its simplicity – there is only one moving part, a spring to push forward the propellant bar. Utilizing

The University of Washington 11 a PPT would negate the need for the expensive and complicated tubing required for a gas feed system.

Solid fuel allows for many benefits related to the system integration, including easy propellant handling and the elimination of valves, feedlines, propellant tanks, and other fluid system hardware. PPTs have a very low dry mass due primarily to the solid fuel. Other advantages stemming from the use of solid fuel include [Peters]:

• No warm-up time required

• Vacuum compatible

• Long shelf life

• Chemically inert/safe fuel

• Minimal temperature requirements and insensitive to rapid changes in temperature

When used solely as a thruster, the pulsed nature is a key advantage to the PPT. Pulsed thrusters can deliver variable thrust at optimum operating conditions by simply adjusting the pulse frequency. In our case this means the ability to adjust the amount of Teflon particles being input to the helicon thruster without opening the chamber. The downside of a pulsed system is that it completely negates the possibility of a DC thruster. Current PPT’s such as on the EO-1 satellite can nominally fire at 1 Hz and demonstrated 20 million successful pulses, which would result in continuous 1 firing every second for over 230 days [Cassady, Zakrzwski]. 1.4 Pulsed Plasma Thruster

A pulsed plasma thruster is a combined electromagnetic and electrothermal thruster that employs a solid Teflon propellant, giving it significant systems-level advantages over other propulsion technologies.

The PPT is a small, self-contained propulsion system that uses solid Teflon propellant, which has the advantage of being inert and non-toxic, giving the PPT system an additional benefit of one of the safest propulsion systems for space travel.

The PPT was one of the first electrical propulsion concepts to see use in space onboard in the Zond-2 satellite in 1964. It is a relatively simple design with high specific impulses, but its low efficiency caused it to be bypassed for more effective and more complex systems in the 1970’s and 80’s [Zakrzwski]. In the past two decades the push for low power, low cost, and high performance propulsion systems has brought a renewed interest in PPT’s.

Presently PPT’s have been used on several recent flight programs and are capable of taking over for several spacecraft applications. Some common PPT applications include attitude control, orbital transformers and maintenance, and de-orbiting.

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The most severe downside to the PPT is its low efficiency, which generally falls in the range of 5-15%. This is due in large part to its poor mass utilization efficiency (< 50%) as a result of two processes – macroparticle ejection and late time ablation [Spanjers]. Not all the ablated propellant is entrained in the current sheet, leading to a significant spread in the exhaust velocity distribution and a large fraction of the total shot mass exiting at a low speed (macroparticle ejection). Furthermore, the thin layer of Teflon heated by the arc continues to evaporate neutral gas for as long as a millisecond after the discharge ends [Spanjers 1996], compared to the tens of µsec which a typical PPT pulse lasts (late time vaporization). Spanjers estimated that this late-time ablation could account for up to 40% of the mass ejected per shot.

1.4.1 PPT Operation

The largest advantage to the PPT is their simplicity. Figure 1.8 shows a simple, breech-fed PPT design with rectangular electrodes. Solid Teflon is used as the propellant and is spring fed to a pair of copper electrodes in a rail configuration. One of the electrodes also has an igniter plug, located at the face of the Teflon bar as illustrated in Figure 1.8. A capacitor is charged to a moderate voltage, run through a transformer that increases the voltage upwards to 16-18 kV to power a spark plug that creates an initial discharge across the surface of the Teflon. This spark, created at the base of the cathode, emits electrons that provide a path for a surface discharge of the capacitor across the Teflon propellant face. A standard-sized PPT will store tens of Joules in the main capacitor and develop a peak current in the kiloamp range. This current pulse ablates several micrograms of material from the propellant surface, generating a Teflon plasma. Flowing upward in the figure, the current creates a self-induced magnetic field orthogonal to the current vector. The interaction between this magnetic field and the charged particles in the plasma, the interaction known as the Lorentz force, drives the current sheet along the electrodes and expels the ablated propellant out of the thruster at a high velocity, generating thrust.

Figure 1.8. Schematic of a typical PPT with rectangular electrodes [Eckman].

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Because there are no pressurized liquids or gasses involved in the thruster, there is no need for valves or storage tanks. The only moving part of the propellant feed system is the spring, which pushes the fuel bar forward as the end is ablated. This combined with the stability and durability of the Teflon propellant makes the PPT an extremely safe and reliable thruster.

PPT’s may be classified into a 2 x 2 matrix of the four types:

Rectangular Coaxial Breech-fed Side-fed

As stated above the most common configuration is the breech-fed PPT with parallel rectangular electrodes, driven by a long current pulse, as typified by the LES-8/9 and EO-1 thrusters. For the rectangular electrode configuration, the current and self-induced magnetic field are primarily transverse to the flow. The anode and cathode arc attachment regions are then free to move downstream along the electrodes with the flow. [Burton and Turchi]

The coaxial configuration is usually conical, typically with central and downstream ring electrodes. In any PPT the igniter plug is usually mounted in the cathode, which can be either in the center electrode or the downstream ring.

Breech-fed PPT’s feature a current distribution that initially propagates downstream at a speed greater than the mass-average velocity, suggesting that the mass rate of ablation at the rear of the arc is not sufficient to provide spatial stabilization. In this design the primary force vector points away from the ablating surface, so that when the arc strikes it moves off the surface, reducing the ablation rate. On the other hand, for side-fed devices, the vector is parallel to the ablating surface, and the magnetic field lines are normal to it, providing a conduit for electrons to reach the Teflon and maintain the ablation rate. Thus, the downstream motion of the side-fed PPT, to first order, does not change the ablation rate [Burton and Turchi].

In addition to the four basic geometric and propellant feed variations, PPT’s can also be distinguished by second order variations, such as propellant type, electrode direction, and pulse shape. Although dozens of solid propellants have been tested, none have produced as high and as reliable performance as Teflon. Teflon provides attributes of high ISP, high impulse bit, and zero surface charring. In the rectangular geometry the plane electrodes are usually parallel, but for some thrusters the performance can be improved by angling the electrodes away from the thrust axis. Although this generally produces an increase in the electromagnetic acceleration of the plasma, the plasma density is reduced by an order of magnitude or more [Burton and Turchi].

Based on the geometry and its application, current day PPT’s have Isp ranges from 300 to 2000 seconds with thrusts from 50 – 2000 µN at energy levels below 50 J. These values are detailed by mission in Table 1.2.

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1.4.2 PPT Plume

The plume of a pulsed plasma thruster consists of numerous ionized and neutral particles of decomposed Teflon. Polytetrafluoroethylene (PTFE), or the more commercial name Teflon, is a synthetic fluoropolymer of tetrafluoroethylene with numerous applications in everyday life [Giner and Hunter]. It is composed of long polymer chains with four fluorine atoms per every two carbons (C2F4). Gaseous decomposition of Teflon in vacuum has been found to generate neutral molecules in the following amounts: 94% C2F4, 2.6% C3F6, 0.86% CF4, and 0.73% C4F8 [Wentink]. A number of other propellants have been researched with hopes of finding improvements over Teflon; however, Teflon is still the primary PPT propellant as it allows for a small, self-contained, inert and stable propellant system.

As discussed in Section 1.4.1, a large current sheet is expelled from the thruster at an extremely high velocity. However, not all the ablated propellant is entrained in the current sheet, leading to a significant spread in the exhaust velocity distribution and a large fraction of the total shot mass exiting at low speed. This is primarily due to two separate inefficiencies: macroparticle production, and late time ablation.

While the electromagnetically accelerated plasma can be ejected at tens of km/s, slower particles exit at only a few hundred m/s [Burton and Turchi]. [Spanjers] captured images of these millimeter-sized macroparticles as they were ejected from the Teflon surface and found them to travel at an estimated 300 m/s, far too slow to contribute to appreciable thrust. Studies of PPT exhaust plumes on the LES 6, LES, 8/9, and EO-1 satellites found ion and electron exhaust speeds ranging from 12 – 40 km/sec [Burton and Turchi, Zakrzwski, Byrne, Eckman]. A study on the USAF Millipound Thruster which used witness plates and collimated quartz crystal microbalances to make plasma measurements concluded that 20-30% of the exhaust plume remained a polymer (neutral) and traveled at velocities below 500 m/sec [Dawbarn].

The second primary inefficiency in PPT performance is late time ablation (LTA). LTA is the sublimation of Teflon that takes place after the main discharge, due to the Teflon being at a temperature above its sublimation point. This thin layer of Teflon heated by the arc continues to evaporate low speed gas and macroparticles for perhaps as long as a millisecond after the discharge ends. For comparison, the typical PPT arcs in this thesis lasted for 50 - 80 microseconds. This so-called late time ablation is hypothesized to account for up to 40% of the mass ejected per shot [Marques], and does not contribute noticeable thrust to the PPT.

The combination of these two inefficiencies leads to only 30-40% of the ablated Teflon being ionized and accelerated to high exhaust velocities [Marques] and is the main factor in limiting the PPT efficiency to 5-15%.

Another limiting factor in PPT performance is related to component lifetimes. Specifically, high-voltage capacitors will ultimately fail after repeated cycling and/or voltage reversals [Rayburn]. This effect is more severe when capacitors are cycled

The University of Washington 15 near their maximum rating, so taking a mass penalty and choosing a capacitor with a large voltage margin can alleviate this difficulty. Spark plugs are frequently used to initiate PPT discharged in a controllable way and tend to limit the thruster lifetime as material is eroded away. It should be noted that these lifetime issues are not insurmountable, as the LES 8/9 PPT successfully fired 20 million pulses [Eckman].

A typical PPT discharge may be modeled as an inductance-resistance-capacitance (LRC) pulse circuit [Jahn]. Since the plasma resistance is very low, this circuit is almost always under-damped, so the current oscillates for a few periods before the capacitor is fully drained. This tendency to oscillate not only strains the capacitor, it can also lead to a phenomenon known as restrike in which a second, less energetic current sheet forms behind the first and ablates more material [Jahn]. This tends to reduce the efficiency since the capacitor is mostly empty at the time of the restrike and because the new conductive path prevents further current from flowing to the original, energetic current sheet. Diodes could be used to reduce the possibility of restrikes [Jahn].

1.4.3 PPT Flight History

Pulsed plasma thrusters have a flight history spanning over the last five decades. The Soviet Union was the first to develop the PPT as an EP concept, designing two PPT designs in 1962, one with an electromagnetic and one with a thermal acceleration mechanism (Figure 1.9). The latter version was launched on the Zond-2 spacecraft for a Martian flyby in 1964 [Burton and Turchi]. Six coaxial breech-fed PPT’s were onboard to provide three-axis attitude control in order to keep the solar arrays facing the sun [Shelton]. Six months after launch radio communication with the satellite was lost, and with it control of the PPT’s. Propulsion characteristics for the Zond-2 PPT are given in Table 1.2.

Figure 1.9. The first PPT designs from the Soviet Union. The electromagnetic (right) and electrothermal (left) designs were both considered. The Zond-2 spacecraft used the electrothermal design [Burton and Turchi].

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The first American PPT was designed at MIT’s Lincoln Laboratories and flown in 1968 to provide East-West stationkeeping capability to the 6th Lincoln Experimental Satellite. The LES-6 PPT performed nominally for the 10-year mission while in orbit [Burton and Turchi].

The rectangular breech-fed design to the LES-6 PPT provided the blueprint for the missions throughout the 1970’s. The United States launched the Synchronous Meteorological Satellite (SMS) in 1974, whose PPT was the first to have variable thrust capability and was used for spin axis precession control [Burton and Turchi]. The Lincoln Laboratories continued research and flight qualified a PPT for use on the LES 8 and 9 satellites, both of which employed a rectangular geometry and had an increase in the specific impulse by a factor of 3 and an increase in the thrust by an order of magnitude [Burton and Turchi] compared to the SMS satellites. However, in a last minute change, the PPT’s were scratched from the mission in favor of a cold-gas propulsion system.

Figure 1.10. Shown are the Soviet Union’s Zond-2 Satellite (left) and the American LES-6 satellite (right) [Burton and Turchi]. Both utilized pulsed plasma thrusters as their primary propulsion system.

The development of PPT’s continued at a modest pace in the 1980’s with the launch of the Navy’s three NOVA navigational satellites in 1981; all of which employed TRANSIT Improvement Program-II (TIP-2) PPT’s for drag makeup to maintain a “drag-free” orbital position [Ebert]. These thrusters saw an increase in the produced impulse bit from 300 µN-sec on the LES 8/9 design PPT to 378 µN-sec, but decreased the overall thrust from 600 to 378 µN and the efficiency from 7 to 3% [Cassday]. The Chinese and Japanese also utilized PPT’s in the early 1980’s onboard their MDT-2A and ETS-IV satellites, respectively. The MDT-2A satellite was an early suborbital test for the Chinese space program, while the ETS-IV used a PPT to provide spin rate control for the Japanese satellite [Burton and Turchi].

Over the following decades PPT research and development accelerated due to a growing interest in smaller, lighter, and longer duration missions. The most recent The University of Washington 17

PPT hardware tested in orbit was onboard the first Earth Observing spacecraft (EO- 1) launched by NASA in 2001. This thruster provided attitude control for the spacecraft without negatively impacting the spacecraft bus or onboard science instruments [Zakrzwski]. The primary improvement over the LES PPT’s was the drastic reduction in the electronics mass to approximately 5 kg. EO-1 is the largest PPT to have ever been flown, producing a thrust of 860 µN (0.0002 lbs) with 9.8% efficiency [Zakrzwski]. This force is equivalent to what a human would feel by holding a 2 by 2-in piece of paper; however in the space environment, where there is no atmospheric drag, this thruster could propel a spacecraft to 30,600 miles per hour (mph) – almost twice that of the space shuttle (18,000 mph) [Zona].

Figure 1.11. The EO-1 spacecraft with attached pulsed plasma thruster (left). The PPT is attached to the right of the spacecraft. In the image on the right, a 12-inch ruler is shown to give scale [Zakrzwski].

In the early 2000’s, the University of Washington, through the Aeronautics and Astronautics department, developed one component of the ION-F (Ionospheric Observation Satellite Formation) nanosatellite program, the Dawgstar, which used a PPT based on a scaled down EO-1 design [Rayburn]. This PPT was designed to be the primary propulsion component for the satellite and can be seen in testing in Figure 1.12. The University of Washington, along with Utah State University and Virginia Polytechnic Institute, each designed a nanosatellite for the mission – which aimed to demonstrate formation flying, miniature bus technologies, and distributed satellite capabilities. Dawgstar was capable of providing a thrust-to-power ratio of 8.8 µN/W and an ISP of 482 sec at only 5.2 J of energy [Rayburn]. Over the last three years the Earth and Space Science department at the University of Washington in conjunction with Eagle Harbor Technologies, have investigated a micro pulsed inductive thruster (µPIT). It is currently being tested as a means to increase PPT propellant utilization while maintaining a low overall mass, so to provide a new alternative to the current generation of micropropulsion technologies. µPIT was capable of providing 5 µN/W with an ISP of 2000 seconds at only 0.4 J of energy while weighing 0.377 kg (combined thruster and PPU mass) [Peters]. The thruster is shown in Figure 1.12. The University of Washington 18

Figure 1.12. The Dawgstar PPT in operation (left) [Rayburn] and the µPIT thruster nozzle (right) [Peters].

Table 1.2. The thrust and impulse characteristics of the pulsed plasma thrusters discussed previously. All the listed thrusters have a firing rate of 1 Hz. For variable thrust capability, the maximum values were listed [Burton and Turchi unless otherwise noted]. * signifies the PPT was never flown.

Mission E Ibit Isp Thrust Spc. Thrust Eff. Purpose (J) (µN-s) (sec) (µN) (µN/W) (%) Zond 2 50 n/a 410 2000 40 n/a Attitude Control LES 6 1.8 25 310 25 10.6 3.9 E/W Station Keeping SMS 8.5 110 450 130 12.2 5.8 Spin Axis Control LES 8/9* 25 300 1000 300 12 6.8 Attitude Control [Eckman] TIP-II / NOVA 20 400 850 375 13.3 3 Orbit Insertion & Drag Correction EO-1 56 860 1396 860 12.3 9.8 Attitude Control [Zakrzwski] Dawgstar* 5.2 55 625 55 8.8 3 Attitude Control [Rayburn] µPIT* [Peters] 0.4 120 2000 n/a 5 18 Micropropulsion

1.5 Outline and Research Goals

The primary goals of this project are to design, build, and test a low-cost PPT. High speed cameras and Langmuir probes were used to experimentally measure the efficiency and effectiveness of the PPT while characterizing its shape and behavior.. Photographs were used as an aid to modify the design to increase and isolate the plasma ionization region, while the Langmuir probes were used to characterize the plume of the PPT. Based on these measurements we were able to calculate a number of thruster performance values, such as the impulse bit, efficiency, and specific impulse. We then investigated the feasibility of installing the PPT onto the HPH system by placing the Helicon magnet system over the PPT in the bell jar to determine how the external magnetic field would affect the plume of the thruster

The University of Washington 19 and how the plume would affect the magnet structure itself. In particular the magnetic field should limit the cross-field transport of electrons between the anode and cathode which could prevent the PPT from firing if the magnetic field was strong enough. A number of integration work left to accomplish was identified and will be completed over the coming years.

Chapter 2 describes the vacuum system, data acquisition, and other infrastructure necessary for the experiment. Chapter 3 describes the diagnostics implemented for the PPT. The theory behind the data reduction techniques are presented in Chapter 4. Chapter 5 gives an account of the multiple iterations of the PPT and the performance results found. Chapter 5 also discusses the results found when pairing the PPT with the Helicon magnets. Finally, Chapter 6 contains a summary of the major contributions of this project and recommendations for the direction of future work.

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2 Experimental Setup and Electronics Design 2.1 Vacuum System

The Advanced Propulsion Laboratory’s bell jar, shown in Figure 2.1, was used for these series of experiments. A mechanical roughing pump drove the pressure down to ~500 millitorr in about 10 minutes, then a turbopump took over, reducing the chamber pressure down to a base value of 9 x 10-7 torr. However, the data presented in this thesis was collected in the range of 1 x 10-6 to 3 x 10-6 torr. The Varian TV 550 Navigator turbopump had a pumping rate of 550 L/s for nitrogen at a maximum speed of 42 k-rpm. The roughing pump is attached in series with the turbopump so that the turbopump exhaust flows directly to the roughing pump inlet. There is no gate-valve on the chamber.

Figure 2.1. The bell jar vacuum chamber enclosed within a plexiglass shield. The mounting flange is bolted to the table and the glass jar is connected to the flange only through the pressure difference imposed upon the o-ring.

The bell jar measures 46 cm in diameter by 75 cm high, which provides ample room to study the evolution of the exhaust plume. The stainless steel mounting flange, also shown in Figure 2.1, adds 23cm in height and provides several access ports to the chamber. Eight one inch radial ports accommodate power leads, input/output data connections, and a venting valve. Not shown are the four ports located at the base of the mounting flange below the table. One of these serves as the inlet for the turbopump while another provides access for the cold cathode and chamber Pirani gauge. The other two provide Langmuir probe access.

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Figure 2.2. Shown here are the turbopump, roughing pump, and all three gauges connected to the chamber. This setup lies directly below the bell jar.

A Pirani gauge measures the pressure in the tubing between the turbopump output and the roughing pump inlet. The Pirani senses pressure by passing a constant current through a filament; the filament temperature increases as the gas pressure drops and a thermocouple senses the temperature directly. The Pirani gauge can accurately measures between 4 millitorr and 5 torr. Another Pirani measures low vacuum in the bell jar itself, while a cold cathode gauge measures high vacuum in the bell jar. The Televac 7E cold cathode gauge can measure pressures between 1 x 10-8 and 1 x 10-2 torr. The cold cathode gauge is a type of ionization gauge that measures pressure by bombarding a gas with electrons and collecting the resulting ion current. With a calibration, this ion current can be related to the gas pressure. 2.2 Control and Data Acquisition

Custom LabView software was used to control the optical pulse generator, which sent signals to the scopes, camera, and the initial arc control circuit in the form of light along optical fibers. However, all of these devices require electrical digital logic signals, so the optical pulses were translated to 5 volt TTL (transistor-transistor logic) signals with nearby custom-made receivers. Pulse width and timing were adjustable from the LabView interface. These optical fibers are widely used in fiber- optic communication instead of metal wires as they permit transmission over longer distances, at higher bandwidths, and signals travel along them with less loss and electromagnetic interference [Bates].

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2.3 Electronic Control Systems

Due to the simplicity of the PPT, the corresponding electronics are inherently straight forward, as shown in Figure 2.3. The starter arc consists of a transformer used to bump up the voltage of the small charging capacitor. This transformer is connected directly to the igniter and cathode. An IGBT switch system controls when the starter fires. The main discharge has a large capacitor strapped across the PPT electrodes in parallel. A diode is placed after the positive lead of the cap to prevent any of the starter high voltage to negatively recharge the cap. The main discharge fires when the starter arc has produced enough seed electrons and ions to form an electrical connection between the anode and cathode.

Figure 2.3. The electronic schematic for the PPT showing both the starter arc and main discharge.

2.3.1 Starter Power Processing Unit

The starter arc utilized a circuit board with 10 Insulated-Gate Bipolar Transistor's (IGBT) (Figure 2.4 only shows five as the other five are on the bottom of the board), a snubber capacitor of 0.15 uF, and a main capacitor bank with five 0.22uF caps in parallel, each cap is rated to 2kV. This gives an effective capacitance of 1.1*10-6 F, and at 500 V an energy of 0.14 J per discharge. The logic side of the circuit board requires 18 V to operate and a fiber optic cable directs the IGBT’s when to open and close. A 76-2 iron-core hand-wrapped transformer was built to increase from the starter voltage from 500 V up to 19 kV. The starter would arc at 16 kV approximately 90% of the time and at 19 kV had a 99% arc rate. The transformer was placed inside the bell jar to reduce the length of the high voltage leads. The output of the circuit board is run through a stanganese, and then directly to the vacuum chambers feed-through, to the transformer, and then to the spark wire. The SP system is powered off a 3.1kV, 30mA power supply. The circuit was dumped through four 40 ohm, 25-watt resistors. This design was chosen to simplify access to the circuit board and to decrease and confine the length of the high voltage wires.

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Figure 2.4. The circuitry for the starter arc which is housed within the electronics box with the main discharge circuitry.

2.3.2 Main Discharge Power Processing Unit

The main discharge circuit consists of a 60 uF K75-40a Russian built capacitor (capable of storing 4 kV) connected in parallel to the PPT electrodes. Two 15 kV diodes were connected in series before the anode to reduce ringing within the circuit and to prevent the high voltage in the igniter circuit to flow back into the 60 uF capacitor. The dump circuit board which was used to safely discharge the main capacitor is shown in Figure 2.5. Two resistors, placed in series, of 250 Ohms each were used to dump the energy of the 60 uF capacitor in less than a second. Each resistor was capable of handling 225 W of power. A 1:100 voltage divider (described in Section 3.2) was also attached in parallel across the capacitor leads that resulted in a RC discharge time of 30 minutes, if the primary dump circuit failed. Ross relays were used to switch the capacitor from charging to discharging. A 10kV, 10mA high voltage power supply was used to power the circuit.

A Stanganese was placed between the anode and cathode posts of the capacitor and was used to measure the current flowing out of the capacitor. The voltage across the capacitor was measured on an oscilloscope by a high voltage probe across the cap during each discharge as well as being continuously read off by a voltmeter connected across the voltage divider.

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Figure 2.5. The main discharge circuitry is shown here, including the 60 μF capacitor, 15 kV diodes and the dump board.

Two 15 kV high voltage international rectifier diodes were installed in series with the high voltage port of the capacitor to prevent the capacitor from receiving a current due to the backflow from the discharge. These can be seen wrapped in protective plastic on the far left of Figure 2.5. Common capacitor and voltage traces for the main discharge are shown in Figures 2.6 and 2.7 with and without the diode installed. The need for the diode can be seen by comparing the current traces without the diode (left) and with the diode (right). Without the diode on the 120 J discharge the current spiked positive at 5820 Amps and then went negative to - 2050 Amps. With the diode installed, the current again spiked at 5820 Amps, but only drops to -100 Amps, where the diode then stops the flow of current. It can also be seen that both the current and voltage traces are extended in shot length with the diode. This is believed to be due to the increase in inductance when connecting the diode to the circuit. The diode also eliminated the unexpected jump in the voltage and current that occurred at 52 µsec for all four energy levels. It is not known what this discontinuity is created from, however it could be an effect from the capacitor when the current became too negative.

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Figure 2.6. Generic voltage traces for the PPT main discharge without the diode (left) and with the diode (right). The traces are shown for four different energy input levels (120 J in blue, 43 J in green, 20 J in red, and 10 J in turquoise).

Figure 2.7. Generic current traces for the PPT main discharge without the diode (left) and with the diode (right). The traces are shown for four different energy input levels (120 J in blue, 43 J in green, 20 J in red, and 10 J in turquoise).

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2.4 HPH Base Magnets

The helicon thruster to which this PPT will eventually be attached requires a base magnetic field to guide the helicon wave propagation. This was provided by a set of six coils, each consisting of 43 turns of copper wire as shown in Figure 2.8 in operation with the PPT and Figure 1.7 in operation with the HPH system. These coils, whose axes are aligned, are spaced 2.2 cm apart, have a radius of 7.3cm, and are 10 cm thick. A 10 kW ECR power supply (40 V, 250 A) provides DC current to the magnets. The magnetic coils can be connected in a variety of configurations, however they are typically operated with all six magnets energized to create a solenoidal magnetic field of 300 G that points downstream. Therefore, that was the configuration which was tested around the PPT. The magnet system was placed in the bell jar over the PPT with the Teflon surface between the 2nd and 3rd magnet from the bottom (approximately where the back end of the helicon antenna would be). These magnets were only placed over the PPT at the end of the experiment and were not used for the majority of the density and temperature measurements as they are not part of a typical PPT.

PPT Electrodes

Figure 2.8. Base magnets over the PPT in the bell jar while at atmospheric pressure. A Langmuir probe can be seen in the right image 20 cm above the Teflon surface.

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3 Diagnostics

Several instruments were used to help characterize the PPT’s electrical systems and exhaust plasma in order to estimate some key performance parameters. Taken together, these instruments allowed for a greater understanding of the physics of the device and guided changes and improvements. This chapter discusses the theory underlying Stangenes probes, voltage dividers, Langmuir probes, and the fast- framing CCD camera. 3.1 Stangenes Probe

Stangenes probes (or current transformers) are used to measure flowing electric currents. Like any other transformer, a stangenes probe has a primary winding, magnetic core, and a secondary winding. The alternating current flowing in the primary produces a magnetic field in the core, which then induces a current in the secondary winding circuit. Transformers of 100-1 and 10-1 were used to characterize the current flowing in the SP and main discharges. 3.2 Voltage Dividers and the High Voltage Probe

Voltage dividers were constructed to monitor the storage capacitor and the voltage applied across the electrodes over the course of each shot. Voltage dividers are attractive because they are simple and inexpensive to implement. They also provide a neutral discharge path for capacitors that could otherwise hold a potentially dangerous charge for an extended period of time. However, they also come with some disadvantages. Voltage dividers require a direct connection to the circuit, applying a high impedance load. In the case of high voltage measurements like the PPT electrodes an arc or a failure can lead to high voltage appearing across the sensor signal leads, endangering the digitizer. Voltage dividers are also highly susceptible to noise. An example of a voltage divider circuit in shown in Figure 3.1.

Figure 3.1. Voltage divider circuit diagram. For this application and .

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The circuit produces an output voltage given by:

(3.1)

Typically the resistances are chosen such that in order to divide down to a manageable signal. This expression is valid as long as nearly all the current flowing through R1 also flows through R2, which can be enforced by choosing an input impedance to measure that is much larger than R2. The voltage divider used here was constructed with ¼ Watt carbon film resistors to measure the charge on the storage capacitor. R1 was set at 900 kΩ and a proportionality constant of 100 was selected, resulting in R2 being 9.1 kΩ. Calculations verified that the peak power dissipation was well below ¼ W and that the RC time was approximately 54 minutes, much greater than the characteristic PPT shot time of 100 µsec.

While Eq. 3.1 is accurate, large resistances can be difficult to measure with precision. For this reason, the voltage divider constructed for the PPT was calibrated using a Tektronix P6015 high voltage probe. This probe was also connected across the capacitor electrodes during operation and its signal was recorded on an oscilloscope. 3.3 Langmuir Probes

Electrostatic, or Langmuir probes were the first diagnostics developed capable of taking measurements inside of a plasma. Physically they are very simple consisting only of exposed conductive wires biased to some potential within a plasma. One of the drawbacks of most Langmuir probes is that a voltage vs. current curve is needed to properly determine plasma properties from the probe measurements. This V-I curve is developed by sweeping the voltage applied to the probes and measuring the resulting current collected; it is used to determine electron temperatures, plasma density, and the floating potential within the plasma. There are a number of different variations of the Langmuir probe. The single probe consists of a single wire protruding into the plasma and a conducting wall in good contact with the plasma in order to complete the circuit. The glass bell jar used to test the PPT does not meet this requirement. An alternative is the double Langmuir probe, which is electrically floating with respect to the chamber wall but maintains a potential difference between the two probe tips. First developed by Johnson and Malter, the double Langmuir probe can provide local density and electron temperature measurements [Johnson and Malter]. Double Langmuir probes were used in this study to map the density profile throughout the PPT plume and to measure the exhaust velocity using a time-of-flight technique. This section outlines the standard Bohm theory for a collisionless, quiescent plasma. It was assumed that effects on the probe due to the flowing plasma were minimal and were therefore ignored.

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3.3.1 Standard Double Probe Theory

Figure 3.2. The circuit diagram for a double Langmuir probe.

A double Langmuir probe consists of two probe tips immersed in the plasma with a voltage applied between them, as shown in Figure 3.2. The probe is floating so that it draws no net current. Figure 3.3 schematically shows the relevant potentials for double probe operation. The Bohm theory relies fundamentally on the formation of a sheath around the probe tips. A sheath naturally forms around any object inserted into the plasma because the highly mobile electrons carry more current to the probe than the ions. If electrically isolated, the probe will quickly charge to a potential such that the flux of ions and electrons results in zero net current. The voltage the probe acquires in this process is the floating potential ( ), and is necessarily more negative than the plasma or space potential ( ).

Figure 3.3. Double Langmuir probe potential diagram [Byrne].

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As long as the probe potential is less than the plasma potential, ions are accelerated and electrons are retarded in a sheath region of a few Debye lengths across. As the figure indicates, the application of the bias voltage ( ) tends to drive one probe above the floating potential and the other below. This asymmetry combined with the requirement for zero net current causes P1 to collect a net electron current while P2 collects a net ion current. Several assumptions are necessary to develop Bohm theory, the most commonly used theory for interpreting double Langmuir probe I-V characteristics. First, a Maxwellian velocity distribution is assumed, meaning the average velocity of the particles is zero as the particles constantly interact with each other, but move freely between short collisions.. Next, the plasma is assumed to be collisionless on the order of the probe radius

(3.2) and on the scale of the sheath thickness

(3.3)

Here, λ is the mean free path and the subscripts refer to collisions among ions, electrons, and neutrals. is the sheath thickness, given approximately by

(3.4)

In the above equation, is the Debye length, where , defined as

(3.5)

and Vo is the potential difference between the probe tip and the plasma. Probe tips should be arranged with enough spacing to prevent sheath interactions. For a probe spacing s, this requirement can be expressed as

(3.6)

It is also assumed in the following analysis that gradients in the collection region are negligible, that no magnetic fields are present, and that the plasma is stationary. This way, the probes only collect random current due to thermal motion. Obviously, the plasma being stationary is a problem. The current to each probe tip ( ) can be expressed in terms of the combined ion and electron current. The convention used here is to treat the collected electron current as positive and the collected ion current as negative.

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(3.7)

The electron current will be addressed first. The electron saturation current density can be expressed in terms of the thermal speed .

(3.8)

The thermal speed for either species is given by

(3.9)

Equation 3.8 gives the electron saturation current density, which is the current density one could expect if the sheath voltage did not interfere with electron collection. A Maxwellian distribution is assumed in order to account for the fact that only the most energetic electrons are collected by the probe; the rest are reflected. Consequently, the electron density in the sheath is reduced by the Boltzmann factor

(3.10)

The electron current to the probe with sheath area As can then be expressed using this Boltzmann factor

(3.11)

Unlike the electrons, the ions are assumed to be cold with an energy approaching zero at infinity. The ions gain energy as they fall through the sheath. Bohm showed that the formation of a sheath with a sheath voltage Vs requires that

(3.12) in order that the ion flux matches up at the interface between the plasma and the sheath. Quasineutrality in the region just outside the sheath requires that the ion density match the electron density given in Eq. 3.10. Additionally, the ion velocity entering the sheath can be approximate by

(3.13) if the electron charge is taken to be positive. Using Eqs. 3.10 and 3.13, an expression for the ion current density can be written

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(3.14)

Eq. 3.14 can then be combined with Eq. 3.12 to write the Bohm formula

(3.15) where the Bohm velocity is given by

(3.16)

Substituting Eqs. 3.10 and 3.15 into 3.7 and assuming that the sheath area equals the probe surface area (As = Ap) yields the current collected by a single probe tip.

(3.17)

The measured current I shown in Fig. 3.2 is the difference between the currents collected by the probe tips

(3.18)

This may be combined with the fact that the probe draws no net current

(3.19)

With these relationships between I1 and I2 and Eq. 3.17, a probe I-V characteristic can be written in terms of the ion saturation current and the probe bias voltage.

(3.20)

Note that the two probe collection areas were assumed to be equal in this analysis and that the sheath areas were assumed to match the probe areas. This latter approximation is most accurate in the thin-sheath limit, in which

(3.21)

Eq. 3.20 is plotted below in Figure 3.4. A noteworthy feature is the way the curve flattens out for large bias voltages. The theory predicts that the current will be independent of the probe potential, a result of the thin sheath approximation and the assumption that the ions are cold. Increasing the probe potential while already

The University of Washington 33 in saturation tends to increase the sheath thickness, resulting in a large collection area and a slowing increasing saturation current. The assumption that the probe areas are equal leads to a symmetric I-V characteristic with positive and negative ion saturation values that are equal in magnitude.

Figure 3.4. Theoretical IV characteristic for a symmetric double Langmuir probe.

3.3.2 Probe Construction and Description

A symmetric double Langmuir probe, shown in Figure 3.5, similar to those used in other PPT experiments was constructed and used to sample the charged particles in the plume. The probe tips were made of 1.1 mm tungsten wire, which were fed through an aluminum tube. Both tips were shaped to remove any sharp points that might radically change the electric field. These wires were epoxied into a ceramic tube and electrically connected via crimps to magnet wire which was fed through a 1/4” stainless steel tube to a quick disconnect out of the bell jar. The grounded steel tube served to reduce electromagnetic and electrostatic noise. The formvar- insulated leads were twisted together to minimize magnetic pickup. The probe tips were 4.6 mm long and presented a cross-section to the flow of 5.06 mm2.

Figure 3.5. The symmetric double Langmuir probe used for these experiments.

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The Langmuir probe circuit (Figure 3.2) was kept as simple as possible. The voltage between the tips was supplied by a bank of batteries to isolate the circuit from electrical noise due to power supplies or contact with the chamber ground. The length of the time the probe saw plasma was typically on the order of a few hundred microseconds at most. Typical batteries are too slow to give up current during this time, so a fast acting (1 µF) capacitor was used to eliminate the problem. The capacitor could handle up to 400 V. The probe was isolated from the oscilloscope via a 1:1 Stangenes isolation current transformer. Typically the current from the probe passed through the isolation transformer 4 times, making it a 4:1 transformer.

Traditionally double Langmuir probes are used on steady plasmas as this allows for the bias voltages to be swept continuously through a range of values while measuring the resulting current flowing through the probe system. From these measurements, the characteristic curve Figure 3.4 can be reconstructed and the electron temperature and density can be found quite easily. However, a PPT discharge produces an unsteady plasma, lasting less than 100 µsec. This transient plasma makes the voltage sweep impossible in this time period. This was overcome by sweeping the bias voltage by discrete levels over the course of several pulses. Shot to shot variations in the thruster were expected and resulted in uncertainties of 15% in the density measurements at 20cm above the Teflon surface and 20% at 40cm above the Teflon surface. Each test point had 4-6 Langmuir measurements taken, and the resulting values averaged together to reduce the shot to shot variation error.

The Langmuir data was imported into the LabVIEW graphing program where the first 300 data points were averaged together to correct any offset and the stangenes effect was corrected. 3.4 CCD Camera

The final diagnostic tool used to study the PPT plume was a CoolSnap CF fast- framing CCD (Charge-coupled device) camera. This camera is controlled with a TTL pulse sent from the LabVIEW control software. The camera can capture images with a minimum exposure time of approximately 10 µsec. The camera was used to capture plume images of each PPT variation experimented with. Since the shot time is much longer that the minimum exposure, several images were captured at different stages during the current pulse in order to gain information about the time evolution of the plume. A delay of approximately 5 µsec exists between the time the camera receives a trigger pulse and the start of an exposure. A small variation in this delay from shot to shot contributed to uncertainty in the portion of the current pulse that was actually imaged. The camera was also used with longer exposure times to capture images of entire shots.

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4 Theoretical Models and Data Reduction

A technique was developed to estimate the thruster efficiency and impulse bit using density profiles, time-of-flight (TOF) velocity results, and capacitor voltage measurements. The plasma temperature was found through Eq. 3.20, which allowed for the density calculation to be made with Eq. 3.15.

It is important to note that the Langmuir probes only measured the electron density. However, while neutral gas is expected to make up the majority of the ablated mass per shot, the low velocity associated with neutrals means that they contribute very little impulse. For example, neutrals evaporating from a Teflon surface at 600K (approximate minimum temperature for Teflon decomposition [Turchi et al]) would have a thermal velocity of about 550 m/s compared with 10-55 km/s for the electromagnetically accelerated propellant in typical PPTs [Burton].

The TOF velocity technique was compared to the calculated thermal velocity (Bohm velocity), which is a typical speed of the thermal motion of the particles which make up a plasma [Peters]. Technically, it is defined at the width of the peak in the Maxwell-Boltzmann particle velocity distribution. It can be mathematically displayed as the root mean square of the velocity in one dimension:

(4.1)

The thermal velocity was compared to the measured flow velocity using the TOF method. The symmetric double Langmuir was positioned at 20, 30, and 40 cm on- axis above the Teflon surface. The separation between the probes was always known to within 2 mm. Density data was collected on the same oscilloscope to ensure synchronized waveforms. The time between the density peaks was measured to an accuracy of approximately 0.5 µsec (500nsec). The known probe separation and time lag between density peaks enabled the calculation of the plume velocity. Based on the accuracy of the time and distance measurements, velocity uncertainty was estimated at 16% for 10cm probe separation and 11% for 20cm probe separation.

With an exhaust velocity know, the specific impulse was calculated. The specific impulse ( ) is a value to describe the efficiency of a rocket or thruster and is related to the thrust as , where is the Earth’s gravitational constant . It represents the impulse (change in momentum) per unit amount of propellant used and can be written as

(4.2)

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By measuring the profile of the plasma plume the mass flow rate of the electromagnetically charged particles could be calculated with

(4.3)

The mass of the charged particles in the plume was found by multiplying the mass flow rate by the time of the shot. An ion mass of 2.5 × 10-26 kg was assumed. This was taken by averaging the ion mass between carbon (6 amu) and fluorine (9 amu). If we assume that only 30% of propellant is ionized [Spanjers, Marques], then the total amount of propellant ablated was able to be calculated. The reasoning for the low ionization rate, macroparticle ejection and late time ablation, is described in Section 1.4. The ablated mass per shot was estimated at each energy level in Section 5.2.7.

The impulse is the force applied by the thruster over a set time. Therefore the impulse bit is defined as the impulse provided per shot and is calculated by

(4.4) where the shot mass is only the mass of the charged particles as the force provided by the neutrals is negligible.

The thrust efficiency is defined as the ratio between the impulse generating energy and input energy.

(4.5)

The input energy can be represented as the energy drained from the storage capacitor per shot length as follows, where is the capacitance, is the initial storage capacitor voltage, and is the final voltage. As the capacitor was found to fully drain on each shot, equals zero.

(4.6)

The output energy in the flow, or the impulse generating energy, is defined as (4.7) which therefore allows the efficiency to be expressed as

(4.8)

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5 PPT Iterations and Discussion

The PPT chosen for this experiment was a parallel-plate PPT, in the heritage of the EO-1 satellite. This geometry was chosen over coaxial electrodes primarily for its longer successful flight heritage.

Initially a PPT of the same size as EO-1 was examined. However, high speed imaging showed that a more contained plasma ablation was required for the HPH purpose. As such a 2nd generation PPT which was roughly 50% smaller was the final design which was examined in detail for this project. 5.1 PPT-1

5.1.1 Spark Plug – PPTv1.0

The first iteration of the PPT (v1.0) is shown in Fig. 5.1 and was based on the mechanical design of the EO-1 satellite. The rectangular electrodes were made from copper and had an exposed length and width of 2.5 cm. A block of Teflon was contained within a housing made of nylon and fed between the electrodes with a spring. The Teflon had a frontal area of 2.5 x 5 cm and was held in placed with a small 3 mm notch in the anode. The cathode had a 2 cm diameter hole cut out and a Champion Aerospace (REM37BY) spark plug imbedded within it, giving a total thruster mass of 1.8 kg. The spark plug was powered from the control systems discussed in Sect. 2.3.1. The control board was placed outside the chamber, but the transformer placed within to shorten the length of the high voltage leads. The main discharge leads were screwed directly into the copper electrodes and the cathode lead was grounded to the chamber walls.

Figure 5.1. The first iteration of the PPT with the spark plug imbedded within the cathode.

There were two main concerns with the spark plug. Firstly it weighed 122 grams, increasing the mass of the thruster by 10%. Secondly the spark plug initially required 17 kV to fire in vacuum and 5 kV to fire in atmosphere. However over time the spark plug begun requiring additional voltage to operate in a vacuum, increasing to 22 kV before not firing at all. It was decided that a simpler design consisting of just a wire arcing to the cathode would be used. This was the basis for the second iteration of the thruster.

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5.1.2 Spark Wire – PPTv1.1 & 1.2

This spark wire entered through a 0.25 cm diameter hole in the cathode before arcing to the cathodes front face as shown in Figure 5.2. The wire was insulated with kapton tape and heat shrink, leaving only the tip exposed to arc. It required 16 kV to fire in vacuum and 4 kV up to air, while weighing only 15 grams with its insulation, effectively solving both concerns over the spark plug.

Figure 5.2. The PPT with a spark wire (v1.1) is shown here at 20 µsec into the pulse (top right), 60 µsec (lower left), and 100 µsec (lower right). The PPT was fired with an input voltage of 300 V.

Based on the images takes with the high-speed camera (Figure 5.2), it was not believed that this design was producing significant quantities of Teflon. Large amounts of ionization were occurring at the anode and a strong “beam” was seen traveling from the spark wire to the top of the anode. Furthermore, light production could be seen at the high voltage connection point on the anode and little, if any light production was seen traveling across the Teflon surface. The spectrometer traces taken did not shown high quantities of carbon or fluorine, but confirmed oxygen molecules in the PPT plume – consistent if blasting off pieces of the copper electrodes.

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PPTv1.2 looked to improve upon the previous designs by addressing the concerns listed above. Torr seal was added on the back and sides of the electrodes in order to decrease the number of field lines produced off the Teflon surface. The effectiveness of this can be seen by comparing Figures 5.2 and 5.3 and noticing the drastic increase in ionization across the Teflon surface, which should indicate an increase in the density being produced.

Figure 5.3. PPTv1.2 is shown here with torr seal added to the back and sides of both electrodes and Kapton tape added over the high voltage screw connected to the anode. The images are shown at input voltages of 200 V (top right), 300 V (lower left), and 400 V (lower right).

Although more ionization was occurring over the Teflon surface, it was hoped that more light would be seen. The EO-1 satellite, which this PPT was modeled after, was designed to not only produce Teflon particles, but also accelerate them up to high speeds. The longer the electrodes, the larger the force on the plasma will be, which should accelerate the particles to a higher velocity. However, smaller electrodes should dump the same amount of energy, only now with more of that energy closer to the Teflon surface, theoretically burning off more particles at a lower velocity. The present PPT only needs to produce Teflon particles, as the helicon coil will provide the majority of the acceleration. Figure 5.3 shows that the majority of ionization across the Teflon surface results from the lower half of the

The University of Washington 40 electrodes. Additionally, the wider the block of Teflon the more particles there are to be ablated off, and theoretically, higher densities are possible. However, as it appeared that the majority of the arc was not traveling directly across the surface of the Teflon, the width of the block was reduced. 5.2 PPT - 2

5.2.1 No Insulation - PPTv2.0

The second version of the PPT had a Teflon block only 2.5 cm across and electrodes 1.3 cm high. No torr seal was initially added and even so, a drastic increase in the amount of ionization occurring across the top of the surface of Teflon was noticed, as shown in Figure 5.4.

Figure 5.4. PPT 2.0 is shown at 20 µsec into the pulse (top left), 40 µsec (top right), 60 µsec (bottom left), and 80 µsec (bottom right) with a voltage of 400 V.

Although improved, this strange beam-like feature was still evident traveling from the spark wire to the top of the anode, and ionization still occurred on the backside of the anode and down near the nylon housing. Insulation was used to correct these inefficiencies in v2.1.

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5.2.2 Insulation - PPTv2.1

Kapton tape and torr seal were both experimented with to reduce the extraneous ionization and the torr seal was found to produce slightly higher densities. Additionally,the Teflon block was further reduced in length to just 1.3 cm across. Figure 5.5 shows that with these modifications, the entirety of the ionization occurs over the Teflon block and between the electrodes, which was the desired result.

Figure 5.5. PPT2.1 is shown at 200 V (1.2 J) and 40 µsec into the shot. The torr seal was placed over the anode only – including the connection point where the high voltage lead was soldered to the electrode.

Increasing the exposure time on the CCD camera to 70 µsec and filming from behind the anode resulted in Figure 5.6. The plume which was lit up due to the ionization appeared to be tilted off-axis at approximately 5 degrees. It is believed that this effect is due to inconsistencies in manufacturing of the electrodes and application of the torr seal.

Figure 5.6. End on views of the PPT firing at 1200 V (43 J) with an exposure covering the entire length of the shot.

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The Langmuir probe described in Section 3.3.2 was used to measure the density and temperature of the plasma. Data was taken from 20 to 40 cm in 5 cm increments, No measurements could be taken below 20 cm due to the Langmuir sputtering effect which is commonly seen in hot plasmas. This effect caused a giant leap in the voltage output which the oscilloscope recorded. An example of this is shown below in Figure 5.7. The traces were both taken at 20cm above the thruster surface; however the large jump seen on the red trace is not believed to be real, but rather due to this sputtering effect.

Figure 5.7. An example of the sputtering affect on Langmuir probes at 43 J. The blue trace was taken at a battery voltage of 120 V, while the red trace was taken at 129 V. In this case the maximum density level and time could be recorded, however if the sputtering effect had occurred before the peak, then the data would be useless.

5.2.3 Electron Density On-axis

Typical density traces taken without sputtering at 20 cm above the Teflon surface on-axis are shown below overlaid with corresponding Gaussian approximations in Figure 5.8. Gaussian approximations have been shown to accurately approximate PPT exhaust plumes to within 85% [Rudolph, et. al.]. The Gaussian approximations used are listed below in Eq. 5.1. As is evident from the Figure, the Gaussian overestimates the density at the beginning of the shot and then underestimates the density at the end of the shot. The traces in Figure 5.8 were used to find the data points of the Langmuir probe in saturation for Figures 5.9-5.11.

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(5.1)

Figure 5.8. Typical, non-sputtering Langmuir traces taken at 10, 20, and 43J at 20cm as a function of time. These three traces correspond to the dark blue traces in Figures 5.9 – 5.11.

Figures 5.9-5.11 and Table 5.1 show the Langmuir current sweep results taken at 10, 20, and 43J in 5 cm increments from 20 to 40 cm. All Langmuir probe measurements have a 15% uncertainty in the density measured. The 10J energy level at 40 cm was too weak to consistently make out and as such as excluded. The density at 20 cm on-axis ranges from 5.6 × 1017 to 2.6 × 1018 m-3 for the 10 J and 43 J energy levels respectively. At 35 cm this decreases to 2.2 × 1017 and 1.3 ×1018 m-3 for the 10 J and 43 J levels respectively. The University of Washington 44

Figure 5.9. Electron density and temperature results on-axis at 10J.

Figure 5.10. Electron density and temperature results on-axis at 20J.

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Figure 5.11. Electron density and temperature results on-axis at 43J.

Table 5.1. The electron density and temperature Langmuir data on-axis for the PPT at 10, 20, and 43 J and heights from 20 to 40 cm in 5 cm increments.

Energy (J) Height (cm) Density (m-3) Temperature (eV) 20 5.6 × 1017 30 25 3.7 × 1017 25.9 10 30 3.0 × 1017 17.8 35 2.2 × 1017 20.6 40 n/a n/a 20 1.3 × 1018 23.0 25 1.1 × 1018 22.4 20 30 7.7 × 1017 21.8 35 6.2 × 1017 21.8 40 5.0 × 1017 21.7 20 2.6 × 1018 47.0 25 1.7 × 1018 35.4 43 30 1.5 × 1018 32.1 35 1.3 × 1018 20.7 40 1.0 × 1018 27.8

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At 20 cm downstream, the PPT on the LES8/9 satellite recorded electron densities of 1.5 × 1020 and 4.2 × 1020 m-3 at 20 and 40 J, respectively [Eckman]. This is roughly two orders of magnitude higher than the density found here. Likewise, [Eckman] found electron temperature to be between 2.5 and 3 eV. This factor of ten decrease in the temperature would correspond to an order of magnitude increase in the density, partially explaining the variation in numbers. Additionally, the LES 8/9 block of Teflon was 50% larger than the block being used here, allowing for more particles to be ablated off. Likewise the EO-1 PPT found similar results with electron densities of 1.2 × 1021 m-3 and a temperature of 2 eV, 20 cm downstream of the PPT at 50 J [Byrne]. The amount of neutrals from the EO-1 PPT and this PPT are unknown, however it can be assumed that a drastically higher temperature should produce more neutral particles. And since the helicon antenna (Section 1.3) will further ionize the Teflon, it is desirable to produce a large neutral population.

5.2.4 Plume Width

Previous PPT experiments have modeled the exhaust plane as a function of radius at set heights with a standard Gaussian function, peaking on the centerline for both coaxial and rectangular PPTs [Zwahlen, Burton and Turchi]. Peak density measurements were taken at 6 and 8 cm off-axis from behind the anode at heights between 20 and 40 cm. The 6 and 8 cm were measured from the edge of the Teflon surface. These measurements allowed a Gaussian approximation to be made of the exhaust plume shape and the width of the plume to be calculated. These measurements are shown in Figures 5.12-5.14, for the 10, 20, and 43 J cases respectively and summarized in Table 5.5. It is important to note that these figures show how the plume expands radially at one instant within the shot. Figure 5.8 on the other hand shows how the density varies over time at one particular location within the shot.

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Figure 5.12. Peak electron density on-axis and at 6 and 8 cm off-axis at 10 J at heights between 20 and 35 cm above the Teflon surface.

Table 5.2. Constants in Eq. 5.2 used for the 10 J Gaussian density approximations in Figure 5.12.

Height (cm) a (m-3) c (cm) 20 5.6 × 1017 7.7 25 3.7 × 1017 9 30 3.0 × 1017 12.5 35 2.2 × 1017 20.1

The Gaussian approximations at 10 J fit the measured data points to within 15% at all four heights. The Gaussians all took the form of Eq. 5.2 with the constants shown in Table 5.2. As is evident from the figure, the plume approximations are denser and narrower at low heights and as the plume rises, the peak density decreases and the width of the plume increases. Clearly there is some uncertainty in these Gaussian approximations as only 3 data are considered for the curve. More measurements taken at distances further from the centerline would help the accuracy of the model; however we were limited by the access ports in the bell jar and concerns over anomalous effects increasing as the sides of the chamber walls were approached.

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(5.2)

The Gaussian approximations at 20 J fit the three data points to within 22% at all five heights. The Gaussians used all took the form of Eq. 5.2 with the constants shown in Table 5.3. Again, as is evident from the figure, the plume approximations are denser and narrower at low heights and as the plume rises, the peak density decreases and the width of the plume increases. The rate at which the plume width increased was calculated with Eq. 5.3 and shown in Figure 5.15.

Figure 5.13. Peak electron density on-axis and at 6 and 8 cm off-axis at 20 J at heights between 20 and 40 cm above the Teflon surface.

Table 5.3. Constants in Eq. 5.2 used for the 20 J Gaussian density approximations in Figure 5.13.

Height (cm) a (m-3) c (cm) 20 1.3 × 1018 8 25 1.1 × 1018 6.4 30 7.7 × 1017 11.8 35 6.2 × 1017 15.3 40 5.0 × 1017 28.5

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The Gaussian approximations at 43 J fit the three data points to within 18% at all five heights. The Gaussians used all took the form of Eq. 5.2 with the constants shown in Table 5.4. The approximations showed the same trends as in the previous two figures.

Figure 5.14. Peak electron density on-axis and at 6 and 8 cm off-axis at 43 J at heights between 20 and 40 cm above the Teflon surface.

Table 5.4. Constants in Eq. 5.2 used for the 43 J Gaussian density approximations in Figure 5.14.

Height (cm) a (m-3) c (cm) 20 2.6 × 1018 13.9 25 1.7 × 1018 21.3 30 1.5 × 1018 18.8 35 1.3 × 1018 21.2 40 1.0 × 1018 30.3

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Table 5.5. The on-axis and off-axis electron density measurements taken at 10, 20, and 43 J. Off-axis measurements were made at 6 and 8 cm from the Teflon edge.

Electron Density (m-3) Energy (J) Height (cm) On-axis 6 cm off-axis 8 cm off-axis 20 5.6 × 1017 4.2 × 1018 3.4 × 1018 25 3.7 × 1017 3.0 × 1018 2.5 × 1018 10 30 3.0 × 1017 2.7 × 1018 2.4 × 1018 35 2.2 × 1017 2.1 × 1018 2.0 × 1018 40 n/a n/a n/a 20 1.3 × 1018 8.5 × 1017 7.8 × 1017 25 1.1 × 1018 7.3 × 1017 5.3 × 1017 20 30 7.7 × 1017 6.7 × 1017 6.2 × 1017 35 6.2 × 1017 5.8 × 1017 5.2 × 1017 40 5.0 × 1017 4.9 × 1017 4.8 × 1017 20 2.6 × 1018 2.4 × 1018 2.2 × 1018 25 1.7 × 1018 1.6 × 1018 1.6 × 1018 43 30 1.5 × 1018 1.5 × 1018 1.4 × 1018 35 1.3 × 1018 1.2 × 1018 1.2 × 1018 40 1.0 × 1018 1.0 × 1018 0.9 × 1018

The previous three figures showed that as the height increased, the width of the plume also increased. For any smooth function, the full width at half maximum (FWHM) is an expression of the width of the function, given by the difference between the two extreme values of the independent variable (radial distance from the centerline) at which the dependent variable (electron density) is equal to half of its maximum value. When applied to a normal Gaussian, such as Eq. 5.2, the FWHM is known to go as

(5.3)

This equation was applied to the Gaussian approximations in Figures 5.12-5.14 and the resulting relationships (Figure 5.15) at 10, 20, and 43 J were found. The linear lines of best fit, when the y-intercept was set to zero, were found to be:

(5.4)

Previous experiments found that 90% of the electrically magnetic particles in the plume of a coaxial PPT are confined to a 40° cone for energies of 5 – 15 J [Rudolph, et. al.]. Here, the FWHMs were found to expand along 47.5, 49.2, and 58.6° cones for energies of 10, 20, and 43 J respectively. The 10 and 20 J cases had similar cone angles, however the much more powerful 43 J case showed a larger cone angle.

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Therefore, it could be concluded that larger energy levels result in larger opening angles for the plume cone, which would be consistent with Rudolph’s work showing that low energy PPT’s have a smaller plume width than found in our case.

Figure 5.15. The Gaussian Approximations Full-Width Half-Maximum values for 10 (green), 20 (red) and 43 J (blue) between 20, 25, 30, 35, and 40 cm above the Teflon surface. Opening angles of 47.5, 49.2, and 58.6° were found for the 10, 20, and 43 J cases by taking the tangent of the slopes of the lines of best fit.

Conservation of flux says that the particle flux under the density curves should remain constant at all heights at any one energy level. Therefore if the integral of the Gaussian density approximation (Eq. 5.2) are evaluated from -∞ to ∞, the resulting electron area density should be constant between heights. This statement makes the approximation that the recombination rate between ions and electrons is small over the range of the Langmuir probe. The integral is shown in Eq. 5.5.

(5.5)

This definite integral was evaluated at each height within each energy level with the results shown in Figure 5.16 and Table 5.6. Average values were taken at each

The University of Washington 52 energy level and found to be 9.9 × 1016, 2.5 × 1017, and 7.9 × 1017 m-2 for the 10, 20, and 43 J energy levels. The calculated area densities are within 16% to their averages for the 10 and 43 J cases and within 30% for the 20 J case. It was expected that there would be some recombination of the charged particles as the height increased, resulting in a decrease in the electron area density with height. The 43 J case shows a slight decrease from 20 cm to 35 cm and both the 10 and 20 J cases show decreases from 20 to 25 cm. However none of the energy levels show consistent area density decreases over the entire height range.

Figure 5.16. The electron area density calculated values based on Eq. 5.5 for 10, 20, and 43 J taken between 20 and 40 cm.

Table 5.6. The peak electron area density average values and their corresponding uncertainties.

Energy (J) 10 20 43 Electron Area Density (m-2) 9.9 × 1016 2.5 × 1017 7.9 × 1017 Uncertainty (m-2) 1.6 × 1016 1.1 × 1017 1.2 × 1017 Uncertainty (%) 15.8 29.7 13.2

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5.2.5 Predicted Electron Densities at 10 cm height

No Langmuir could be gathered on-axis at 10 cm due to the sputtering effect detailed in Figure 5.7, however density measurements were taken at 6 and 8 cm radially off axis (Table 5.7). Assuming that the area densities calculated in Figure 5.15 hold at 10 cm, the off-axis data points should be able to predict the density on axis at 10 cm.

Table 5.7. The off-axis electron density measurements taken at a height of 10 cm at 10, 20, and 43 J energy levels.

Measured Electron Density Calculated Electron Calculated Density Energy (m-3) Density (m-3) Uncertainties (m-3) (J) 6 cm off-axis 8 cm off-axis On-axis 10 4.0 × 1018 2.6 × 1018 7.0 × 1017 2.4 × 1017 20 9.5 × 1017 6.7 × 1017 1.7 × 1018 9.8 × 1017 43 1.0 × 1018 9.6 × 1017 1.0 × 1019 3.1 × 1018

These measurements, combined with the corresponding average area densities in Table 5.6, allowed a Gaussian of the form of Eq. 5.2 to be fit for each energy level at 10 cm above the Teflon surface. This method resulted in calculated peak on-axis electron densities of 7.0 × 1017, 1.7 × 1018, and 1.0 × 1019 m-3 for the 10, 20, and 43 J cases respectively. These peak densities were the a values for Eq. 5.2 and the c coefficients were 5.7, 5.8, and 3.2 cm for the 10, 20, and 43 J cases respectively. Using Eq. 5.3, FWHM’s of 13.4, 13.7, and 7.4 were found for the 10, 20, and 43 J at a height of 10 cm. These widths, if placed on Figure 5.15 at 10 cm, would all fall within the errorbars of their corresponding energy levels line of best fit.

5.2.6 Time-of-Flight Velocity Data

Section 4 details the time of flight (TOF) method for finding the plasmas velocity. The results from the TOF method are shown below in Figures 5.17-5.19, for the 10, 20, and 43 J cases, respectively. As discussed in Section 4, when comparing the results for probes 10 cm apart the uncertainty was 16%, while for probes 20 cm apart the uncertainty was 11%. In order to sync the density results, the current and voltage traces were corrected so to start at 10 µsec. The density results were corrected according to their corresponding current trace and thus insured that traces from different shots could be compared. The additional uncertainty in Figures 5.17-5.19 due to the spread of the data presented was solely due to different scales being selected on the oscilloscope and not due to any physical effects. The trace at 40 cm for the 10 J energy level was an anomaly in the data set and had trouble being repeated, which is why it was not included in Figure 5.9 or 5.12. The Langmuir probe had difficulty detecting such low densities.

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Figure 5.17. TOF velocity data from the electron density results at 10 J.

Figure 5.18. TOF velocity data from the electron density results at 20 J.

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Figure 5.19. TOF velocity data from the electron density results at 43 J.

Table 5.8. A summary of the TOF velocity measurements made at 10, 20, and 43 J in km/s and eV. Uncertainty in Langmuir measurements give errorbars of ±6 km/sec and ±14 eV.

Energy Height Peak Calculated Velocity (J) (cm) Time 20 - 30 30 – 40 20 – 40 (µsec) cm cm cm 20 42 20 km/s 20 km/s 20 km/s 10 30 47 62 eV 62 eV 62 eV 40 52 20 31 25 km/s 10 km/s 14 km/s 20 30 35 97 eV 15 eV 31 eV 40 45 20 32 33 km/s 10 km/s 20 km/s 43 30 35 170 eV 15 eV 62 eV 40 42

At 10 J, density peaks were found at 42, 47, and 52 µsec for heights of 20, 30, and 40 cm respectively. This results in velocities of 20 km/sec when comparing all probe positions (20 – 30 cm, 30 – 40 cm, and 20 – 40 cm). At 20 J velocities of 25, 10, and 14.2 km/sec were found and at 43 J, velocities of 33, 10, and 20 km/sec were found. These numbers compare to within 40% to the calculated thermal velocity of 13.8

The University of Washington 56 km/sec, assuming 30 eV and a Teflon plasma mass of 15 amu. When examining the velocities in terms of electron volts, we find that the 10 and 14 km/s correspond to reasonable temperatures of 15 and 31 eV, respectively. However, the higher speeds of 25 and 33 km/s, give extremely high temperatures of 97 and 170 eV, respectively. Previous PPT experiments [Keidar, Eckman] have found that different ion species travel at different velocities in the plume. In this case however, it is believed that the differing velocities are due to inaccuracies in measuring relating the Langmuir Data.

5.2.7 Mass Ablation Estimates

With measurements made about the plume width, velocity, and shot length, the amount of Teflon produced was estimated for each energy level. Eq. 4.3 was used to calculate the charged particle mass flow rate after integrating the Gaussian density profiles to find an average density at 20, 30, and 40 cm for each of the three energy levels. The electron and ion number densities were assumed to be the same and the electron mass was assumed to be negligible. This mass flow rate was multiplied by the shot length to find the mass of the charged particles ablated with each pulse. After using the standard 30% ionization rate [Spanjers, Marques], the neutral mass of particles was estimated, which allowed for the calculation of the total mass ablated and the specific ablation rate (SAR). These results are summarized in Table 5.9. There were discrepancies between the measured masses at varying heights within each energy level, however these differences could be attributed to the imprecise measurements in the Langmuir data as previously discussed. No matter which height was analyzed, it was found that the PPT at 10 J has a SAR less than half that of the 43 J PPT. This was an unexpected result, however in general, other factors such as propellant surface nonuniformities and the roughness of the electrode surface may lead to discharge nonuniformity [Keidar].

Table 5.9. Comparison of the ablated mass per pulse for the PPT at 20, 30, and 40 cm heights for the 10, 20, and 43 J pulse energies.

Energy Height Charged mass Neutral mass Specific ablation (J) (cm) (µg) (µg) rate (µg/J) 10 20 0.9 2.9 0.4 30 1.1 3.6 0.5 40 n/a n/a n/a 20 20 1.6 5.3 0.3 30 2.6 8.7 0.6 40 5.5 18.4 1.2 43 20 8.9 29.9 0.9 30 10.2 34.2 1.0 40 10.3 34.4 1.0

These masses compare well to published values for commercial thrusters (Table 5.10). As was the case with the density and velocity, our PPT has a similar specific ablation rate of the LES 8/9 PPT at 1.3 µg/J, while performing below modern day commercial products, µPPT at 2.4 µg/J and the Illinois electrothermal PPT at 3 µg/J .

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Table 5.10. Comparison of the ablated mass per pulse for the LES8/9 PPT [Vondra, et. al], µPPT at the Air Force Research Laboratory [Spanjers, 2001], and an electrothermal pulsed plasma thruster currently being tested at the University of Illinois [Bushman and Burton].

Thruster Energy Total ablation Specific ablation (J) rate (µg) rate (µg/J) LES 8/9 20 26 1.3 AFRL µPPT 2.5 6 2.4 Electrothermal PPT 10 30 3

This comparison shows that although the charged particle density may be lower, the amount of propellant being ablated by our PPT is on par or possibly greater than a number of commercial products. And as such it can be assumed that more neutral particles are being ablated, which is the point of the experiment.

5.2.8 PPT Performance Estimates

The PPT had a Bohm velocity of 13.8 km/sec which we will assume to be the exhaust velocity as it is a more conservative estimate than the TOF velocity results. The specific impulse is defined as the exhaust velocity over the gravitation constant (Eq. 4.7), resulting in an ISP of 1413 seconds. When compared to Table. 1.2, this ISP is on order with the EO-1 PPT (1396 sec) and larger than the LES 8/9 PPT by 40%. Although the PPT may be producing fewer charged particles than the EO-1 system, it would appear that they have similar exhaust velocities and therefore similar specific impulses.

10 µg of charged particles are assumed to be ablated with each pulse at an energy level of 43 J (Table 5.9). Eq. 4.4 then predicts an impulse bit of 138 µN-sec, which when compared to Table 1.2 is well below the impulse bit produced by the PPT onboard EO-1 (860 µN-sec), and on par with the impulse bit on the SMS satellite (100 µN-sec) and the Dawgstar PPT built at the University of Washington (55 µN- sec).

Per Eq. 4.6 the energy dumped into the system was 43 Joules. With the impulse bit calculated the impulse generating energy (output energy) was calculated with Eq. 4.7 and found to be 1.9 Joules. This results in a thrust efficiency of 4.4% from Eq. 4.8. This value is half the efficiency of the EO-1 PPT (9.8%) but on par with the Dawgstar PPT (3%), another student built PPT at the University of Washington. Although high efficiency is generally, for this application a high neutral population was desired and this low efficiency implies that to be true. It is believed that the EO-1 PPT had a higher ionization rate than is found here and ran at a lower temperature (1-2 eV, instead of 10 – 30 eV), resulting in more charged particles with the EO-1 PPT and more neutrals with this PPT. This should partially explain the decreased efficiency between this PPT and the commercial products.

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5.2.9 Effect of an Applied Magnetic Field

In order to determine if the PPT designed can be integrated with the helicon thruster system, the magnets which confine the helicon wave were placed in the bell jar and operated at the same time as the PPT discharge. As described in Section 2.4 the magnets produce an applied magnetic field of approximately 300 Gauss along the centerline, where the PPT was placed. The Teflon surface of the thruster was placed between the 2nd and 3rd magnetic coil, meaning that there is 21 cm between the Teflon surface and the top of the magnetic housing.

Figure 5.20 shows the PPT in operation with the helicon magnets. When comparing this image to Figure 5.6, it is seen that the area of ionization appears to be larger, filling up the entire interior of the magnet housing.

Figure 5.20. The PPT in operation with the helicon magnets producing a magnetic field of approximately 300 G.

After proving the PPT would still fire within a magnetic field, the quartz tube which will hold the helicon antenna (Figure 1.7) was placed over the PPT. The tube is 7 cm in diameter and 15 cm long and was placed with the bottom aligned with the top of the Teflon block. As Figure 5.21 shows, the plume travels along the inside of the quartz tube (as expected), but with what appears to be a large amount of plasma being knocked off-axis near the thruster exit. This is probably due to the interaction between the bottom of the quartz tube and the PPT plume. The two images in Figure 5.21 have different colors simply due to differences in the exposure time.

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Figure 5.21. The PPT in operation within a 300 G field and a 7 cm quartz tube placed over the thruster exit. The image on the left was taken with a 20µsec exposure length and the image of the right with a 60 µsec length.

The Langmuir probe was placed on-axis, 30 cm above the Teflon surface, and 9 cm above the top of the uppermost coil. The Langmuir data taken at 10, 20, and 43 J with the magnets on and quartz tube in place is shown in Figures 5.22-5.24 by the dashed curves. The Langmuir data taken previously in Figures 5.9-5.11 without the applied magnetic field is shown in comparison by the solid line. The individual data points collected at the discrete battery voltages tested are shown by stars (no magnets) and dots (magnets on).

The density decreases by 20% in the 10 J case from 3.0 to 2.4 × 1017 m -3. The decrease is 32% in the 20 J case, from 7.7 to 5.3 × 1017 m-3, and in the 43 J case the decrease is 28%, from 1.5 to 1.0 × 1018 m-3. As with all the Langmuir measurements made with this setup, the uncertainty level was 15% for every measurement (described in Section 3.3.2). The temperature on the other hand increased in each case by 10%, 14%, and 12% for the 10, 20, and 43 J cases, respectively.

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Figure 5.22. Electron density results on-axis in the presence of a 300 G applied magnetic field at 10J and 30cm above the Teflon surface.

Figure 5.23. Electron density results on-axis in the presence of a 300 G applied magnetic field at 20J and 30cm above the Teflon surface.

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Figure 5.24. Electron density results on-axis in the presence of a 300 G applied magnetic field at 43J and 30cm above the Teflon surface.

As was expected, the density decreased while in the presence of a magnetic field. The magnets should limit the cross-field line transport of electrons between the anode and cathode, which, if strong enough would prohibit the PPT from arcing. It is believed that the density decrease is due to this cross-field limitation of the electron flow and if the magnets were turned up, a larger density decrease would be found until the PPT stopped working altogether. However, clearly the PPT works within the range that the helicon magnets would operate at without a drastic decrease in the density.

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6 Summary

Pulsed plasma thrusters have several inherent advantages derived from their pulsed operation and solid fuel propellant source. The highlights of these include simple construction, extremely low dry mass, and the ability to vary thrust at a constant specific impulse by simply changing the pulse rate. These are all attributes which make the PPT a desirable solid fuel plasma source for the helicon thruster. The low efficiency of the PPT due to the generation of a large neutral particle population with each shot is a significant drawback. Very little of these neutral particles can be ionized and accelerated by the arc, which is what limits the efficiency. However, when used in conjunction with the helicon antenna, which has proven its ability to ionize more particles than other plasma creation courses, this may not be such a drawback [Ziemba].

The specific goals of this thesis were to design, build, and test a pulsed plasma thruster which could be integrated with the helicon system in the Advanced Propulsion Laboratory at the University of Washington and then initially determine the feasibility of combining the two systems. The testing of the PPT was done with a double Langmuir probe capable of measuring electron density and temperature.

A rectangular PPT similar to one onboard the EO-1 satellite was initially used, but this was soon modified as with this particular application a high velocity was not required, but rather ablating off numerous particles was the desired effect. The final PPT was roughly 50% smaller in electrode and Teflon sizing than the EO-1 satellite and produced two orders of magnitude less electrons. It is believed that this density decrease may be due to a higher neutral particle population based on the fact that the efficiency is only 1.8% compared to the 10% for the EO-1 PPT. As the helicon antenna will further ionize neutral particles, this lack of charged particles is not a drawback.

It is recommended that future work on this experiment focus on reducing the PPT mass, finding a improved method to integrate the quartz tube over the thruster, and look to increase the inductance between the main discharge capacitor and the electrodes in order to further increase the neutral population being created.

The PPT was overdesigned in terms of how much Teflon it could hold. The block of Teflon could be reduced by a factor of 4-5 and still serve our purposes for testing the helicon antenna. Reducing the Teflon block would reduce the nylon housing and the mass of the overall system.

Figures 5.20 and 5.21 both show a large amount of undirected ionization around the base of the quartz tube. This corresponds to a loss of particles entering the helicon antenna. Creating a customized quartz tube to fit over the PPT would decrease this loss and could increase the efficiency of the entire system.

As the helicon antenna should be the main ionization source within the system, having the PPT produce more neutrals and fewer charged particles would be a

The University of Washington 63 benefit for the overall system efficiency (although lowing the PPT efficiency). One method would be to elongate the voltage input to the main discharge by increasing the inductance within the electrical system. A lower maximum voltage entering the electrodes for a longer time period should increase the neutral particle density while decreasing the charged particle density. Along with this concept a method of precisely measuring the neutral population would be required. There are a number of ways available to indirectly measure the neutral population, two of which will be attempted after the PPT has been integrated with the Helicon system. Using spectroscopy, the intensity of the neutral Carbon and Fluorine emission lines would give a qualitative idea of the neutral population. Additionally, the gate valve on the main chamber in the Advanced Propulsion Lab can be shut to prevent pumping on the chamber. The pressure gauge should then show an increase in the overall pressure after a recorded number of shots from which we should be able to back out the total number of neutrals produced.

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